MSFC-M'AN-206 ALABAMA N7_-29276 Unclas G3/31 16&70 SKYLAB SATU RN I B FLIG HT MAN UAL' , SEPTEMBER 30, 1972 REPRODUCED BY U.S. DEPARTMENT OF COMMERCE NATIONAL TECHNICAL INFORMATION SERVICE SPRINGFIELD, VA 22161 C-H 20152
Feb 20, 2016
MSFC-M'AN-206
ALABAMA
N7_-29276
Unclas
G3/31 16&70
SKYLAB
SATU RN IB
FLIG HT MAN UAL' ,
SEPTEMBER 30, 1972
REPRODUCED BYU.S. DEPARTMENT OF COMMERCE
NATIONAL TECHNICALINFORMATION SERVICESPRINGFIELD, VA 22161
C-H 20152
MSFC-MAN-206
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•IHI:
NATIONAl- A_'RONAUTICS AND
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------"_i_:::
SPACE
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ADMINISTRATION
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SEPTEMBER 30, 1972
LISTOFEFFECTIVEPAGES
TOTAL NUMBER OF PAGES IN THIS PUBLICATION IS _.7_., CONSISTING OF THE FOLLOWING:
Page IssueNo.
Title Original
A Original
i thru iv Original
1-1 thru 1-13 Original
2-1 thru 2-18 Original
3-1 thru 3-18 Original
4-I thru 4-64 Original
5-1 thru 5-72 Original
6-1 thru 6-33 Original
7-1 thru 7'-24 Original
8-1 thru 8-13 Original
9-1 thru 9-6 Original
A-1 thru A-5 Original
September 30, 1972
/Sb
TABLEOF CONTENTS
SECTION
SECTION
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APPENDIX
000O0O0@00
Page
General Description ........................................................................................ 1-1
Performance .......................................... .......................................................... 2-1
Emergency Detection and Procedures ........................................................ 3-1
S-IB Stage ........................................................................................................ 4-1
S-IVB Stage .................................................................................................... 5-1
Instrument Unit ................................................................................................ 6-1
Ground Support Interface .............................................................................. 7-1
Mission Control Monitoring ............................................................................ 8-1
Mission Variables and Constraints ................................................................ 9-1
Abbreviations, Signs, and Symbols .............................................................. A-1
FOREWORD
This Saturn IB Flight Manual provides launch vehicle systems
descriptions and predicted performance data for the Skylab mis-sions. Vehicle SL-2 (SA-206) is the baseline for this manual; but,
as a result of the great similarity, the material is representativeof SL-3 and SL-4 launch vehicles, also.
The Flight Manual is not a control document but is intended
primarily as an aid to astronauts who are training for Skylabmissions. In order to provide a comprehensive reference for that
purpose, the manual also contains descriptions of the ground sup-
port interfaces, prelaunch operations, and emergency procedures.Mission variables and constraints are summarized, and mission
control monitoring and data flow during launch preparation and
flight are discussed.
This manual was prepared under the direction of the Saturn Pro-
gram Engineering Office, PM-SAT-E, Marshall Space FlightCenter, Alabama 35812.
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SECTION I mmm_
TABLE OF CONTENTS
Skylab Program .............................................................. 1-1Vehicle Profile ................................................................ 1-1
Saturn History ................................................................ 1-4
Design Ground Rules .................................................... 1-6Saturn IB Production ...................................................... 1-8
Range Safety .................................................................. 1-9
SKYLAB PROGRAM.
The two-stage, liquid-propellant Saturn IB launch vehicle (figure
1-1) is utilized in the Skylab program to transport the three-man
crews to the Saturn Workshop (SWS) in earth orbit.
Objectives of the Skylab program are to establish an experimental
laboratory in earth orbit and to conduct medical, scientific, and
solar astronomy experiments. The laboratory, SWS, includes an
orbital workshop (OWS), which is a modified S-IVB stage: a mul-
tiple docking adapter (MDA): an airlock module (AM): an apollo
telescope mount (ATM): and a payload shroud, which is jettisoned
during launch.
The SWS will be launched and inserted into orbit (figure 1-2) from
LC-39A at KSC as part of vehicle SL-I, which utilizes a two-stageSaturn V launch vehicle. Approximately 24 hours later, a manned
command and service module (CSM) will be boosted into a rendez-
vous phasing orbit from LC-39B aboard vehicle SL-2, which utilizes
two-stage Saturn IB launch vehicle SA-206. After rendezvousing,
and docking with the SWS the crew will activate and inhabit the
SWS for a period of up to 28 days and then return to earth viathe CSM. On Skylab missions SL-3 and SL-4, launch vehicles
SA-207 and SA-208 will place CSM's into orbit for revisitation
of the SWS for up to 56 days on each mission.
After a pitch and roll maneuver, initiated 10 sec after liftoff, thelaunch vehicle will fly a time-tih program that provides near zero
angle of attack through the high dynamic pressure region of flightuntil about 2 rain 10 sec after liftoff. This attitude will be maintained
through S-IB/S-IVB separation at 2 rain 22.9 sec after liftoff, until
iterative guidance is initiated at 2 rain 50 sec after liftoff. During
this time period, the expended ullage rockets are jettisoned at 2min 34.9 sec. The launch escape tower will be jettisoned by astro-
naut command about 23.5 sec after S-IB/S-IVB separation. In the
iterative guidance mode the S-IVB pitch and yaw attitude com-mands are issued to obtain an optimal path (that which requires
the least propellant consumption) to achieve the desired end condi-
tions of flight. Guidance cutoff signal is predicted to occur at 9min 50.1 sec after liftoff.
The S-IVB/I U/CSM will be inserted into an 81 by 120 NM elliptical
orbit inclined 50 deg to the equator, at 9 min 60.1 sec after liftoff.
The CSM will separate from the S-IVB at six min after orbit
insertion. At 15 rain after insertion, S-IVB tank venting will com-
mence to passivate the S-IVB. Pressure sphere sating will be ini-
tiated at approximately 1 hr 23 min after insertion.
At approximately 36 min after orbital insertion for SA-206. the
S-IVB/IU will begin maneuvers to maintain a sun-vehicle orienta-
tion in support of Skylab experiment M415 for the remainder ofthe mission. The SA-208 S-IVB/IU will perform attitude maneuvers
during orbit in support of Skylab experiment S150.
See figure 1-3 for a detailed sequence of events.
VEHICLE PROFILE.
Figure 1-4 shows a cutaway profile of the Saturn IB vehicle andidentifies the first powered stage (S-IB), the second powered stage
(S-IVB), the instrument unit (IU), and the major features of these
stages.
S-IB STAGE.
The S-IB stage is an uprated Saturn l series booster manufactured
by Chrysler Corporation Space Division (CCSD). The basic design
concept incorporates Jupiter and Redstone components because
of their high reliability and qualification status. The S-IB stage
is analogous to the R&D S-I stage but has lightened structure,
uprated engines, a simplified propulsion system, and reduced in-strumentat,on.
The main stage body is a cluster of nine propellant tanks. Thecluster consists of four fuel tanks and four lox tanks arranged
alternately around a larger center lox tank. Each tank has anti-sloshbaffles to minimize propellant turbulence in flight. Stage electrical
and instrumentation equipment is located in the forward and aft
skirts of the fuel tanks.
A tail unit assembly supports the aft tank cluster and provides
a mounting surface for the engines. Eight fin assemblies supportthe vehicle on the launcher and improve the aerodynamic charac-
teristics of the vehicle. A stainless steel honeycomb heat shield
encloses the aft tail unit to protect against the engine exhausts.
A firewall above the engines separates the propellant tanks from
the engine compartment. Eight H-I Rocketdyne engines boost the
vehicle during the first phase of powered flight. The four inboard
engines are stationary and the four outboard engines gimbal for
flight control. Two hydraulic actuators position each outboard
engine on signal from the inertial guidance system.
A spider beam unit secures the forward tank cluster and attachesthe S-IB stage to the S-IVB aft interstage. Seal plates cover the
spider beam to provide an aft closure for the S-IVB stage engine
compartment.
S-IVB STAGE.
B
The S-1VB stage is manufactured by McDonnell Douglas As-tronautics Co. (MDAC). It uses a single propellant tank with com-
mon-bulkhead design and is powered by one J-2 engine. The aft
interstage connects the S-IVB skirt to the S-IB spider beam unit.The aft skirt/aft interstage junction is the separation plane.
A closed loop hydraulic system gimbals the J-2 engine for pitch
and yaw control during flight. An auxiliary propulsion system
(APS), using two APS modules on the exterior aft skirt, providesvehicle roll control during flight and three-axis control during the
coast mode. The exact propellant mass load needed for orbitalinsertion with minimum residuals at cutoff is determined before
launch. A propellant utilization (PU) system helps load this accuratemass.
1-1
Section I General D_scription
INSTRUMENT UNIT.
The instrument unit (IU) is a three-segment ring structure manu-
factured by International Business Machines (IBM) Corporation.
It is sandwiched between the S-IVB stage and the spacecraft LM
adapter (SLA). The S-IVB tank dome actually extends into the
IU ring. The unit is an unpressurized compartment with honey-
comb-panel cold plates mounted around the inside periphery to
accommodated the stage equipment. These panels are thermally
conditioned by a stage-oriented system, which also conditions
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C-H 20149
Section I General DAscription
SATURN IB SEQUENCEOF EVENTS
TIME BEFORE LIROFF (MINUTES)
-3:00 -2:00 -1:00
i,i ITERMINAL COUNTDOWN SEQUENCE START
0:00 1:00
S-IB STAGE--
hLAUNCH COMMIT
L LI FTOFF
PREFLIGHT COMPARTMENT PURGE
JEL BUBBLING
FUEL I_ANK PREPRESSURIZING
REPLENISH TO 100% ILOX BUBBLING
LO)_ TANK PRESSURIZING
S-IB AUX HYDRAULIC PUMPS ON
S-IB HEATERS ON
LOX DOME PURGE ON
LPGG LOX INJECTOR PURGE ON
THRUST CHAMBER FUEL INJECTOR PURGE ON
CALORIMETER PURGE ON
• POWER TRANSFER
:_& iGNITION CMD
& IGNITION ENG 5 & 7
& IGNITION ENGINE 6& R
,i IGNITION ENG 2 & 4
d IGNITION ENG I & 3
_ HOLDDOWN AIP,MS RELEASED
TAIL MASTS R_RACTED
SERVICE ARMS RETRACTED
TIME AFTER LIFTOFF (MINUTES)
2:00 3:00 7:00 8:00 9:00 10_00
*BEP"AT'JN_J "" IS.,VBCUTOFE,I,-+--I .....S-IB/S-IVB SEPARATION SEQUENCE
OUTBOARD ENGINE CUTOFF II I I I ULLAGERKTFIR'NG_. SEPARATION CO MMA_t_D "I SEPARATION STRUCTURE SEVERING
RETR_ ROCK',ETS FIRING
_]1_1111_ S-IVB ROLL CONTROL SYSTEM ACTIVELH2RE_:_RcuLATIO_PUMPON
I_J_l LOX RECIR(_ULATI6N PUMP ON
A S-'IIVB EN: GI N E iTART_COM/_AN D
-1.0 0 1.0 +2.0 +3.0 +4.0
TIME FROM SEPARATION COMMAND (SECONDS)
ml SINGLE ENG NE CUTOFF CAPABILITY
MULTIPLE ENGINE CUTOFF CAPABILITY
• i i iPROPELLANT LEVEL SENSORS ENABLED• INBOARD EINGINES C_JTOFF
• l I =OUTBC)ARDIENGINE CUTOFF ENABLED
I I _ OUTBOARD ENGINES CUTOFF
-4- J_ .___ SEPARATION CMD ,_1.- -- S-IVB STAGE .............
PREFLIGHT FWD & AFT COMPARTMENT PURGE1
l II_ LOX TANK PRESSURIZING l
I J LH2 TANK PRESSURIZING I• POWER TRANSFER
J_-2 THRUST CHAMBEI_ CHILLDOWN J j
LOX_,LH2REC_RCULATIONPUMPSON I II I _ ENGINECHLLDOWN
_" _ ENGINE _G NITIO N_
J " jFLT CONTL COMPUTER S-IVB BURN MODE
JJlJJJ_lJ, LO)_TANK FLT PRESS S_S ON j
J J I LH2 TANK PRESS CONTL SW ENABLED
J I •ULLAGE...... ROI • '_ "^CKET'S JETTISON I
I ENG,NEMIXTUREAT__, I I I I
1 PREVALVESOPENJ AUXII_IARY HYDRAULIC PUMP ON II I ,_ROPELLA,_DEPL_IONCUT6PPAR_6_i' '.I I I I I E_G,'EC_OPP*iI I I LOXNPV"L.2LATC.REUEPVALVESOPENI'I'l
N'_ .......... I FZHTCOTD;TTo ;J l l _ I--I -F- ---[--[ ;I- POWERT_NS_B J / / I
• G=ANCERELEASEI I I / / I III =I'NPLG.TCOND,T,ON;NGj II ,,_ K,. I i
-3:00 -2:00 -1:00 0:00 1:00 2:00 3:00 7:00 8:00 9:00 10:00
NOTES
[_SEE H-1 ENGINE SYSTEM
J_SEE J-2 ENGINE SYSTEM
LEGEND
• INDICATES INSTANTANEOUS EVENT
INDICATES OPERATION STARTS STARTS/lI INDICATES OPERATION ENDS
INDICATES OPERATION BEGINS BEFORE INDICATES OPERATION CONTINUES
START OF THE TERMI NAL COUNTDOWN
SEQUENCER
Figure 1-3
C-H 20150
1-3
Section I General Description
SATURNIB CUTAWAY
BOOST PROTECTI
SPACECRAFT
LM A
LH TANKFO_WAf
S-IVB STAq
COMMON
FUEL FEED
S-,BSTAGE_I I
CENTER LOX TANK_
PROPELLANT
SUCTION
HYDRAULIC
HEAT SHIELD
1-4Figure 1-4
MOTOR
MODULE
ENGINES
IMENT
UNIT UMBILICAL
UMBILICAL
SPHERE
TUNNEL
AFT
UMBILICAL
;E ROCKET
TERSTAGE
J-2 ENGINE
BEAM
UMBILICAL
TANK
B AFT
UMBILICAL
HOLDDOWN FITTING
_H-I ENGINE
C-H 14227
equipment in the S-IVB stage forward skirt because of the proxim-
ity. The IU houses equipment that guides, controls, and monitors
vehicle performance from prelaunch operation to the end of activelifetime in orbit.
PAYLOAD.
The payload is a modular stack attached to the launch vehicle
above the IU stage. It consists structurally of a spacecraft lunar
module adapter (SLA), a service module (SM), a command module
(CM), and a launch escape system (LES).
SATURN HISTORY.
Figure 1-5 compares the three launch vehicles now in the Saturn
family. The following narrative traces the historical events leading
to development of the Saturn IB launch vehicle.
PROJECT HISTORY.
In April 1957, members of the Army Ballistic Missile Agency
(ABMA) initiated studies to establish possible vehicle configurations
to launch a payload of 20,000 to 40,000 lbm for orbital missions,
and 6,000 to 12,000 Ibm for escape missions.
By July 1958, representatives of the Advanced Research Projects
Agency (ARPA) showed interest in a clustered booster that would
achieve 1.5 Mlbf thrust with available engine hardware. ARPA
formally initiated the development program by issuing ARPA Order
14-59 on August 15, 1958. The immediate goal was to demonstrate
the feasibility of the engine clustering concept with a full scale
captive test firing using Rocketdyne H-I engines and available
propellant containers. In September 1958, ARPA extended the
program to include four flight tests of the booster. ARPA Order
47-59, dated December 11, 1958, requested that the Army Ordnance
Missile Command (AOMC) design, construct, and modify the
ABMA captive test tower and associated facilities for booster devel-
opment, and determine design criteria for suitable launch facilities.
In November 1958, ARPA approved the development of a clustered
booster to serve as the first stage of a multi-stage carrier vehicle
capable of performing advanced space missions. The project was
unofficially known as Juno V until, on February 3, t959, an ARPAmemorandum made the name "'Saturn" official.
Because of the presidentat order proposing transfer of the Develop-
ment Operations Division of ABMA to the National Aeronautics
and Space Administration (NASA), an interim agreement was
reached between ARPA, NASA, and the Department of Defense
on November 25, 1959. The agreement provided for transfer of
technical direction of the Saturn program to NASA and for reten-
tion of administrative direction by ARPA. ABMA officially trans-
ferred responsibility for the Saturn program to NASA on July I.1960.
On July 28. 1960, the Douglas Aircraft Company was awarded
a contract to develop and fabricate the second stage (S-IV) of the
recommended configuration. The original design concept specified
four Pratt & Whitney 17.5 klbf liquid hydrogen/liquid oxygen
engines (LR-119). The design was later modified to utilize six Pratt
& Whitney 15 klbf liquid hydrogen/liquid oxygen engines(RLIOA-3).
This S-IV stage flew in the final six Saturn 1 flight tests. Enlarging
and refining of this stage design and replacing the six engines with
a single Rocketdyne J-2 engine produced the S-IVB stage used
in the Saturn IB launch vehicle. The basic S-IVB stage described
here also serves as the third stage of the Saturn V launcn vehicle.
but for Saturn V it has engine restart capability, larger APS modules
Section General Description
SATURN CONFIGURATIONS
WEIGHT SATURN IB SKYLAB APOLLOSATURN V SATURN V
DRY 159,000 LB 585,000 LB 553,000 LB
1,296,000 LB 6,221,000 LB 6,495,000 LBLIRTOFF
TMSFC
IMARTIN
MARIETTA
McDONNELL
DOUGLAS
IBM
-- 346 RT
ATM
MDA
AIRLOCK AND
PAYLOAD
SHROUD
INSTRUMENT
UNIT
INORTH
AMERICAN
ROCKWELL
GRUMMAN
I BM
365RT
APOLLO COMMAND
& SERVICE MODULES
LUNAR MODULE &
SPACECRAR LM
ADAPTER
INSTRUMENT UNIT
McDONNELL
DOUGLAS
OWS McDONNELL
DOUGLAS
S-IVB STAGE
NORTH
AMERICAN
ROCKWELL
IBM
McDONNELL
DOUGLAS
CHRYSLER
SATURN I B
2 STAGE
NORTH
AMERICAN
ROCKWELL
APOLLO COMMAND
& SERVICE MODULES
SPACECRAFT
LM ADAPTER
INSTRUMENT UNIT
S-IVB STAGE
S-IB STAGE
S-II STAGE
NORTH
AMERICAN
ROCKWELLS-If STAGE
,/I IllBOEING - STAGE BOE G S-IC STAGE
SKYLAB/SATURN V APOLLO/SATURN V
2 STAGE 3 STAGE
C.H 14547-2
Figure 1-51-5
Section I General Description
with ullage thrust capability, and a flared aft interstage to mate
with the S-II stage.
The first eight booster stages (S-I) were initially designed, devel-
oped, manufactured, and tested by MSFC personnel. Responsibility
for subsequent stages was transferred to Chrysler Corporation Space
Division (CCSD). On August 20, 1963 CCSD was awarded a con-
tract for 14 booster stages, which would be built at the Michoud
plant, New Orleans, Louisiana. Two S-I stages were produced for
the last two R&D vehicles of the Saturn I program, the remaining
12 stages were to be fabricated to the Saturn IB configuration.
The first four Saturn I flights (SAd through SA-4) had no in-strument unit; instead, these vehicles had instrument cannisters.
The next three flights (SA-5 through SA-7) had a pressurized
compartment that provided a conditioned environment for naviga-
tion, guidance, and control equipment. The unpressurized prototype
used on the Saturn IB vehicles first appeared on the SA-8 through
SA-10 Saturn I flights. New equipment packaging techniques per-
mitted each assembly to be pressurized individually as necessary
and eliminated the need for pressurizing the entire instrument unit.
The instrument unit was designed and developed by MSFC person-
nel. In the Saturn IB program, responsibility for the IU was gradu-
ally transferred to IBM. Work on the first four flight models (S-IU-
201 through S-IU-204) was the responsibility of MSFC, with IBM
doing the actual assembly and testing at its Huntsville, Alabama
facility. The S-IU-205 instrument unit was the first produced en-
tirely under IBM responsibility.
CONFIGURATION HISTORY.
Ten R&D vehicles (SA-I through SA-10) constitute the Saturn I
project. Three R&D vehicles (SA-201 through SA-203) and nine
operational vehicles (SA-204 through SA-212) constitute the Saturn
IB program.
The Saturn I launch vehicles were of two basic configurations. Block
I (SA-I through SA-4) consisted of a live booster stage (S-I) and
dummy upper stages (S-IVD and S-VD): and, Block II (SA-5
through SA-10) consisted of an S-l stage, an S-IV stage, and aninstrument unit. Each Saturn IB launch vehicle consists of an S-IB
stage, an S-IVB stage, and an instrument unit. In the Apollo pro-
gram when the launch vehicle combined with the Apollo payload,
its configuration was designated Apollo-Saturn (AS); hence, AS-204and AS-205.
Vehicles SA-206, -207, -208, and -209 have been assigned to the
Skylab program. As a result they are designated SL-2. -3. and
-4, respectively (SA-209 is the backup vehicle designated SL-R).
DESIGN GROUND RULES.
Saturn IB design has one philosophical ground rule: missionachievement with a safe crew-even under the most adverse flight
conditions. This concept reflects in vehicle subsystem design, quality
control, structural safety factors, and performance reserves.
STANDARDIZED NOTATION.
When designing a complex structure like the Saturn IB launchvehicle, accurate location of vertical levels is made possible by
establishing a datum to which all vertical dimensions are referenced:The measurements in inches from this datum are called stations
(figure I-I). Likewise. structural areas are given a reference desig-nation as shown in figure 1-6. This designator, when used on
equipment, indicates its location in the vehicle.
III
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i
'",.,,.u.,.,o.,, I V-
425 420 421
401 12 ENGINE4018 406 AFT INTERSTAGE INTERNAL
406 THRL(ST STRUCTURE406 406 404 AFT SKIRT INTERNAL
406 MAIN TUNNEL407 406 LOX TANK INTERNAL
406 COMJ_N BULKHEAO
416_ _417 406 FUEL TANK INTERNAL
// _" 406 FUEL TANK EXTERNAL410 FORWARO DOME EXTERHAL
414 415 411 FORWARD SKIRT INTERNAL
414 APE MODULE I (PUS I)
415 APE EK)DULE U (P_ III)....... / k _ 427 lie ULLAGE ROCKET. 30" I TO II
424 "_"_ . s_... / _----404 417 ULLAGE ROCKET. 30" Ill TO ,t
/ 419 AFT INTERSTAGE EXTERNAL
L rvl 403 420 RETRKT. E5° TOI t ' 421 RETRKT 45°IITOILI
,. _-\ 422 RET RKTIN$° IU TO tV
# "_\ 424 AFT DOME EXTERNALt 421 425 AUXILIARY TUNNEL
420 .... "_ 426 FORWARO SKIRT EXTERNAL
427 AFT SKIRT EXTERNAL
406
S,IRSTAGE
UNIT IS --
/ /
Fill
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-- UNIT )l
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Figure 1-6
S-lB STAGE UNIT LOCATIOm
UNIT I ENGINE d COMPANTNENTUNIT 2 ENGINE 12 COMPARTMENT
UNIT 3 +UNGINE.ICUNPARTMENTUNIT i ENGINE 14 C_IFKATNENTUNiT 5 ENGINE _ CUNPARTMENTUNiT E ENGINE _ COMPARTIENT
UNIT 7 ENGINE 17 COMPARTMENTUNIT 8 ENGINE 4 CUNPARTMENTUNiT E TUNU_ FNAtlE AREA
UNIT 10 BETMEEN PROPEU.ANT TANKSUNiT 11 ABOVEUPPER TANK $ttLKHEAUNIT OC (UXIUIZER)UNIT Ol THRU _4 (OXI01ZEN)
UNIT FI TXRU F((FUEL)UNIT Ii INSTN COMPT. F2UNIT 13 INETN. COMPT. Fl
UNIT _ FIN dUNIT ll FIN iEUNIT IE FIN eSUNIT )i FIN .4UNIT _ FIN .$UNIT 21 FIN '_
UNIT 22 FIN 17UNIT 23 TIN 4
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C*H 14475
1-6
Section I General Description
OVERALL DESIGN.
Specific ground rules establish the baselines for each particular
vehicle mission. However, certain overall design requirements have
been upheld throughout the Saturn IB design program. These
requirements were levied to assure the success of all launch missions
assigned to the vehicle, and are summarized in the following para-
graphs.
Reliability.
The launch vehicle must be reliable through the complete flight
phase. Reliability performance must be greater than or equal toa 0.88 factor.
Emergency Detection System.
An emergency detection system (EDS) network must be provided
that can instantly detect vehicle failure conditions, and either initi-ate automatic abort or warn the crew that a manual abort is
necessary. Manual abort must be initiated on at least two separateand distinct indications. In the event of conflicting information
between onboard crew displays and telemetry data relayed to the
ground, the onboard information shall always take precedence.
Abort decisions shall be made by the flight crew whenever time
permits: abort shall be manual rather than automatic when possible.
Triple "Redundancy. Triple-redundant circuits with 2-of-3 voting
logic must be used for all automatic-abort signals. Redundantcircuitry must be used for manual-abort indications from the launch
vehicle to the spacecraft.
Single-Point Failure. A single-point electrical failure in the onboard
crew safety system cannot result in abort, neither can a single-point
failure in the EDS circuitry cause a true or false failure from being
detected. All electrical failure possibilities that jeopardize crew
safety must be designed out. Keep the circuitry simple, and usea minimum number of sensors. The object is to sense effect rather
than cause of failure when possible.
Reliability Goals. As a design goal, the probability of detecting
a failure is 0.9973: whereas, the probability of not detecting a false
failure is 0.9997,
VEHICLECOORDINATESYSTEMS
Separation Systems.
Physical separation of the S-IVB stage from the S-IB stage must
be accomplished with retromotors located on the S-IVB aft inter-
stage, using the short coast mode in a single plane. Successful
separation shall occur even should the following conditions combine
(in a reasonable statistical combination):
a. Engine thrust decay variation
b. Engine dynamic thrust vector deviations
c. Retro and ullage motor static and dynamic misalignment
d. Aerodynamic disturbances
e. Single retro and ullage "motor out" condition
Launch vehicle control will not be lost during stage separation
when influenced by maximum limits on angle-of-attack, pitch rate,
attitude angle, and dynamic pressure.
Winds,
a. Structural design will assure a free standing capability in 99.9
percent probability non-directional ground winds and associated
gusts.
b. The vehicle will be capable of launch in 95 percent probability
non-directional ground winds and associated gusts.
c. The vehicle will be capable of flight in 95 percent probability
non-directional _yinds, plus 99 percent associated wind shears andgusts. _._,- _
Structure.
The general vehicle structure will have a yield safety factor of 1.10
and an ultimate safety factor of 1.40. The erected vehicle will be
supported by eight holddown-and-support arms secured on the
launcher pad during all prelaunch operations. The vehicle structure
must arrest and discharge lightning without damage to the vehicle
system. All exterior protrusions will be minimized.
Propulsion.
Propellants and combustibles will not vent into closed compart-
ments. During prelaunch all hydrogen will duct to adequate disposal
systems. The launch vehicle will be held down following ignition
signal to provide sufficient time (approximately 3.0 sec) for the
engines to reach mainstage thrust. The propellant system design
will permit the running engines to consume all propellant in the
tanks should one booster engine fail during flight. Ullage motors
on the S-IVB stage will settle propellants prior to J-2 engine ignition.
Guidance and Control.
Guidance system equipment will accommodate the preset time-tih
program during S-IB powered flight, and will accommodate path-
adaptive guidance during S-IVB powered flight. Refer to the coor-
dinate system illustrated in figure 1-7; this system is described more
fully in Section VI (Navigation, Guidance, and Control). The launch
vehicle flight control computer in the IU will provide control signals
to each stage.
Electrical Systems.
Each stage will have an independent electrical power system. Elec-
SPACECRAFT
COORDI NATES
+Z
+X
/
Nots
THE CREW IS ORIENTED
WITH HEADS POINTED
ALONG THE -Z SPACE-
CRAFT COORDINATE AND
THEREFORE GOES INTO
ORBIT WITH HEADS DOWN.
°Z
/-X
I
+Z
IV
Figure 1-7
LAUNCH VEHICLE
COORDINATES
C-H 14549
1-7
Section I General Description
trical distributors will centrally distribute signals and power to
minimize cable complexity. Simplicity of design (minimum compo-
nents), operation, maintenance, and construction shall be prime
considerations. Modular components will be used whenever possi-ble. All astrionics systems components will be isolated from their
chassis or case, which will bond to a unipotential structure for
electro-magnetic interference elimination. Total checkout of all
components must be accomplished while the vehicle is on the pad.
This checkout procedure shall test pad components also, withoutinterrupting vehicle circuits.
Instrumentation.
Each stage must have integral measuring, signal conditioning,
telemetering, and RF subsystems. The entire system must be inde-
pendent of other electrical systems. No low-level signals (mV range)
will be brought through the umbilical to ground electrical support
equipment. All measuring signals transmitted to the ground must
first be preconditioned to the 0- to 5- Vdc range. Instrumentation
will provide sufficient flexibility to accommodate vehicle-to-vehiclechanges. Measurements shall be limited to the minimum number
required to operate the launch vehicle and to monitor and analyzesuccess or failure.
Environmental Control.
All vehicle interstages and compartments requiring preflight envi-ronmental control will be conditioned from a source external to
the vehicle. The conditioning medium will be changed to gaseous
nitrogen 20 min before loading Iox on the S-IVB stage. This proce-dure reduces the oxygen content to 4 percent (by volume) or less
in the conditioned compartments. All inflight conditioning require-
ments will be stage oriented: however, the IU stage ECS shallalso control environment in the S-IVB forward skirt.
Ordnance Systems.
Every vehicle ordnance system will have a dual ignition power
source. The vehicle destruct ordnance shall be activated by explod-ing bridgewire firing units (see Range Safety contained in this
section). Dual linear shaped charges will sever each fuel and oxi-
dizer tank required to effectively disperse propellants.
Orbital Coast Period.
The combined IU and S-IVB stages will provide attitude stabiliza-
tion while attached to the spacecraft. The stage combination shall
be capable of sustaining attitude control up to 7 hr and 30 rain.
SATURN IB PRODUCTION.
The S-IB, S-IVB, and instrument unit stages of the Saturn IB vehicle
are manufactured in Louisiana, California, and Alabama, respect-ively. The S-IB is static fired at MSFC, and the S-IVB is staticfired at the Sacramento Test Center in California. The individual
stages are transported to KSC for launch as an integrated vehicle.
Figure I-8 presents the various sites at which production of theSaturn IB launch vehicle is conducted.
The S-IB stage is manufactured and assembled at the Michoud
Assembly Facility near New Orleans, Louisiana. The S-IB is trans-
ferred by river barge from the Michoud facility to MSFC at Hunts-
ville for static testing and is then returned to Michoud for poststatic checkout. From there the stage is transferred by river-ocean
barge to KSC.
The S-IVB stage is manufactured and assembled at HuntingtonBeach, California. It is transferred to the Sacramento Test Center
for static firing and checkout: then, to KSC by air.
The instrument unit is fabricated and assembled at Huntsville,
Alabama. After checkout at the Huntsville facility it is packaged
in special environmental containers and transferred by air or waterto KSC.
STAGEPRODUCTIONAND TRANSPORTATI(IN
S-IVB
SACTO (MDAC TEST)
1-8
HUNTINGTON
BEACH
(MDAC MPG)
LEGEND
SYMBOL STAGE TRANSPORTATION DESTINATION
\RIVER BARGE MIC HOUD/MSFC/MIC HOU DS-IB RIVER-OCEAN BARGE MICHOUD/KSC
..... S-IVB GUPPY H B/SACTO/KSC
.......... IU GUPPY MSFC/KSC
Figure 1-8
MSFC HUNTSVILLE
(S-IB STATIC TEST
IBM MFG)
INSTRUMENT UNIT
L__ S-IB
C-H 14513-2
RANGE SAFETY.
The Saturn IB launch vehicle range safety system (figure 1-9)
enables the range safety officer at Air Force Eastern Test Range
(AFETR) to intentionally destroy the vehicle if it should deviate
beyond the acceptable limits of the intended trajectory, or if an
explosion were imminent during the boost phase of powered flight.
The S-IB stage and S-IVB stage each contain an independent range
safety system. Each system consists of redundant secure range safety
command systems and a propellant dispersion system (PDS). The
secure range safety command systems consist of receiving, decoding,
and control equipment. Upon receipt of command signals from
range safety, these systems provide electrical outputs to shut down
the engines and to detonate the PDS ordnance. A built-in timedelay of 4 sec between range commanded engine shutdown and
propellant dispersion provides crew escape time with the launch
escape system (LES) during first stage flight. During second stage
flight, after LES jettison, the range has agreed not to send the
launch vehicle dispersion command after an abort providing thelaunch vehicle engines have terminated thrust. This allows adequate
time for crew escape during abort using the Service Module propul-
sion system. The PDS shaped charges rupture the propellant tanks,
allowing the propellants to disperse and burn, rather than to ex-
plode. The burning propellant results in only a fractional amountof the theoretical yield if the vehicle should explode.
SECURE RANGE SAFETY COMMAND SYSTEMS.
The secure range safety command systems used on the S-IB and
S-IVB stages (figure 1-10) consist basically of the same type of
equipment. Each stage system uses redundant receiving antennas.power dividers, command receivers, digital decoders, and con-
trollers-all connected in parallel. The command systems receive
power from separate batteries in each stage to increase the overallreliability of the systems. The receiving antennas are located on
opposite sides of each stage to insure the reception of range safetycommands, regardless of vehicle orientation to the transmitting
station. All S-IB stage command-system components are installed
in instrument compartment no. 1, except the four receiving antennas
which are panel-mounted in pairs on opposite sides of the stage
at positions I and 111. Only two receiving antennas are used on
the S-IVB stage. These antennas are mounted on opposite sides
of the forward skirt assembly between positions 1 and 11 and
positions I11 and IV. The remainder of the S-IVB stage commandsvstem is mounted on thermo-conditioning panels in the forward
skirt assembly. The panel at position 14 contains the power dividers.
command receivers, and the controllers. The panel at position 16
contains the digital decoders. The thermo-conditioning panels pro-
vide a heat sink to keep the electrical components cool during
flight. The secure range safety command systems in each stage
perform the same basic operation. The receiving antennas couple
command signals from the ground-based range safety transmitter
through the power dividers to the command receivers. The S-1B
stage uses a power divider with each pair of antennas. The output
from each power divider is directed through a directional coupler
to a command receiver. The S-IVB stage uses a hybrid power divider
to which both antennas are connected. This power divider splits
the inputs into eclual strength signals and applies them through
a directional power divider to the command receivers.
Command Receiver.
Each command receiver operates at 450 MHz and demodulates
and anaplifies the range safety command signals for application
to the digital decoder. The receiver consists of a preselector assem-
bly. first IF amplifier, IF bandpass filter, second IF amplifier.
limiter-discriminator, audio amplifier, isolated outputs, high-level
Section I General Description
telemetry output assembly, and RFI filter-voltage regulator. Thereceiver has two isolated audio outputs, but only one is coupled
to the decoder. Two telemetry outputs from the receiver (high-level
signal strength and low-level signal strength) provide measurement
of the RF input signal strength. Only the low-level telemetry signals
from command receivers are fed into the stage telemetry systems.
During flight an RF carrier is continuously transmitted to the
command system. The low-level output from each receiver is then
telemetered back to ground receiving stations where the signal
strength is monitored on auxiliary pad safety panel. Signal strengthmeasurements received from previous Saturn IB flights indicated
adequate signal strength throughout flight and that the command
system would have performed satisfactorily if needed. A 28-Vdc
output from the command receiver provides a power on signal
to the measuring system as soon as power is turned on. The receivers
are turned on during countdown and remain on until the safe
command after J-2 engine cutoff.
Digital Decoder.
The digital decoder provides timing and gating functions, which
determine the validity of received range safety commands. The
decoder rejects command messages containing erroneous signal
characters or erroneous character sequencing, thereby providing
security against enemy intervention and unintentional interrogation
of the secure range safety command system during flight. A
21-character high alphabet derived from seven basic audio-
frequency tone symbols provides a wide range of message formatsthat can be transmitted as command messages to the secure range
safety command system. Each basic symbol is an audio-frequency
tone in the 7.35 kHz to 13.65 kHz range. Simultaneous transmission
of two of the seven basic symbols creates a high-alphabet character.
Eleven high-alphabet characters contained in two words constitute
a command message. An address word uses nine characters to
condition the decoder for receipt of a function word. The function
word contains two characters that produce a decoder output. Be-
cause only nine of the available 21 characters are used in a com-
mand message, the function word uses two of the same characters
as the address word: however, a given character is used only once
in a given word. Identical code plugs installed on the vehicle digital
decoders and on the digital encoder at Range Safety Central Con-
trol, Cape Kennedy, Florida, determine the command message
format. Since only nine characters are used in the address and
function words, the code plug channels the 12 unused characters
to decoder circuitry, which will cause rejection of a command
message if any of the unused characters appear at the decoder
input. Additionally, if any of the correct characters forming the
address or function word arrive out of sequence at the decoder
input, the decoder will reject the command message. Range Safety
Central Control transmits a frame consisting of ten identical com-
mand messages to insure reception by the vehicle command system.
Each character period including dead time is approximately 8.6
msec except the eleventh character period, which is 25.71 msec.
The extended period for the eleventh character insures receipt of
the function word. Each command message has a total time periodof 1 I 1.43 msec.
Controller.
The range safety controller assemblies receive and transfer the
decoder outputs to the EBW firing units and also provide all power
switching for the command system. All measurement signals except
the receiver low-level strength signals are routed through the con-
troller assemblies. The controller assemblies consist of relays, a
resistor diode module, and four electrical receptacles: three for
electrical cable connection and one for installation of a no-sating
plug for S-IB stage controllers or a sating plug for S-IVB stagecontrollers. During prelaunch operations, ground power supplied
through the controller assemblies operates the command system
1-9
Section I General Description
RANGESAFETYSYSTEM
RSCR NO. 2
A18
)NT ASSY
NO. 2 A19
RANGE SAFETY
COMMAND ANTENNA
SAFETY
COMMAND ANTENNAS
PANEL NO. 14
DEVICE 11A12 A11A97
_SSY
NO. 1 A13
HYDB_ D POWER
DIVIDER A15
CH-14367.2
1-10 Figure 1-9
Section I General Description
;ECURERANGESAFETYCOMMAND ;YSTEMS-IB STAGE
POS I
DtRECTIONAL
COUPLER
NO. I__ °
POS POS
POS III I II
\\_/ SAFETY
POWER.'1"IDIVIDER "__o_ _, o
S-IVB STAGE POS POS
III IV
F
SECURE RANGE SAFETY J
COMMAND SYSTEM NO. t J
28 VDC J_ _COMMAND
, ,_ _,-i_,'_ R.'27R
I t_.IIL..T_ r-_oo,
I 2_VDC _ ._?_
I _-_"_'k _ SAFING/I _ I I__ NO-SAE_NG
I '0°7| 28 VDC I L _ If
_ II CONTROLLER_kI_I NO.,-_ hl11_l.d_ -
_±_
COMMAND
RECEIVER NO, 1
k
DIVIDERI_ I _HECKOUT
¢ _ _° _' o , _I'_IGNAL
COMMAND
RECEIVER NO. 2
J
S-IVB STAGE COMMAND SYSTEM
SAME AS S-IB COMMAND SYSTEM
EXCEPT AS NOTED.
*SECURE RANGE SAFETYCOMMAND SYSTEM NO. 2
DECODER_1_NO.2--___
S._'-_NGG"_
28vDcATTE NO.2--EBW FIRING
__ UNTNO2_2300 V
S-IVB STAGE COMMAND SYSTEM USES A SAFING
PLUG. S-IB STAGE COMMAND SYSTEM USES A
NO-SAFING PLUG.
TIME DELAY FUNCTION ACCOMPLISHED BY TIMER
AT CAPE KENNEDY CENTRAL CONTROL. 4-SEC
DELAY FROM TRANSMISSION OF 'ARM' COMMAND
TO TRANSMISSIO N OF 'DESTRUCT' COMMAND.
E_EONATOR EO'ffONATOR
III_1_ SAFETYANDPDS ARMINGDEVIC,
Figure 1-10
C-H 14196-1
Section I General Description
until the power transfer command energizes a magnetic-latching
relay that switches the command system to internal power. If the
command system must effect propellant dispersion during flight,
two command messages must be transmitted to the secure range
safety command systems. The first command message will cause
the decoder to energize a magnetic-latching relay that applies 28
Vdc to the EBW firing unit charge circuitry and 28 Vdc to the
engine cutoff circuitry. The second command message, which arrives
approximately 4 sec later, causes the decoder to produce an output
that triggers the EBW firing unit discharge circuits. The trigger
is coupled to the firing unit through normally closed contacts of
a propellant dispersion inhibit relay in the controller. This relay
stays energized (with the normally closed contacts open) by ground
power until the umbilicals disconnect at liftoffto prevent destruction
of the vehicle on the launcher. After normal J-2 engine cutoff,
the RSO safes the secure range safety command system since the
propellant dispersion capability is no longer necessary. The safe
command causes the decoder, through circuitry inthe sating plug,
to reset the range safety command receiver internal-external powertransfer relay to the external position. This removes the 28-Vdc
power from the decoder and receiver, rendering the command
system inoperable. Once the system has been safed during flight
it cannot be reactivated. The S-IB stage secure range safety com-mand system cannot be safed since no-sating plugs are installed
on the controllers. After staging, the S-IB propellant dispersion
system presents no danger to the S-IVB stage and payload.
PREFLIGHT OPERATIONS.
During preflight operations, test cables connected to the directional
power divider in the S-IVB stage and to the power divider between
the directional couplers in the S-IB stage permit closed-loop check-
out of the secure range safety command systems by using signals
transmitted by range safety. This eliminates the necessity for openly
transmitting the secure code-of-the-mission command signals, which
could seriously compromise mission security. The S-IB and S-IVB
propellant dispersion panels control the checkout of the secure
range safety command systems. Switches control power application,
internal and external, to the command systems. Indicators provide
visual monitoring of the command system condition. Receipt of
the first range safety system command charges the firing units and
issues the engine cutoff signal, which is monitored by the CUTOFF
indicators. Receipt of the second command signal, which triggers
the EBW firing unit discharge circuits is monitored by the PRO-PELLANT DISPERSION indicators. The RSCR 21 SIGNAL
BLOCKED and RSCR #2 SIGNAL BLOCKED indicators moni-
tor the propellant dispersion inhibit relays to insure that the EBW
firing units cannot be triggered during countdown. During checkout
operations the inhibit circuit is disabled by relays in the ML S-IB
and S-IVB program distributors that are patched into the simulateliftoff circuitry. Placing the simulate liftoff switches on the S-IB
and S-tVB propellant dispersion panels in the SIMULATE position
energizes the patched relays, which remove power from the inhibit
relays in the vehicle controllers. This permits the decoder output
to trigger the EBW firing units to check the output pulse. The
firing unit output pulse discharges into a pulse sensor that provides
a FIRED indication on the propellant dispersion panels. After
checkout completion, the pulse sensors are removed and the patched
relays are disconnected to energize the propellant dispersion inhibit
relays in the controllers. Approximately 60 min before liftoff the
secure range safety command systems are switched to internal
power. In the S-IB stage, battery DI0 supplies power to command
system no. 1 and D20 supplies power to command system no. 2.
S-IVB stage bus +4D30 supplies power to the S-IVB command
system no. I and +4D20 supplies power to S-IVB command system
no. 2. The internal power command plus safety-and-arming device
armed signals (Section IV, Ordnance) provide the S-IB and the
S-IVB ORD OK indications on the pad safety supervisor panel.
The range safety carrier being transmitted to the command receiver
is monitored for signal strength as a result of the low-level outputs
from the receivers to the RSCR AGC indicators on the auxiliary
pad safety panel. These indicators are also monitored throughoutthe flight.
GROUND COMMAND STATIONS.
A range safety command transmission may be necessary at any
time from liftoff until after S-IVB stage J-2 engine cutoff, just prior
to orbital insertion. Tracking stations at Cape Kennedy and at
subsequent locations downrange provide vehicle position indica-
tions to the range safety officer (RSO). When the vehicle is below
the radio horizon of Cape Kennedy, the RSO switches all com-
mands from the Cape Kennedy stations to the downrange stations.
The range safety equipment incorporates a priority-interrupt
scheme that will interrupt any command being transmitted by the
AN/FRW-2A transmitter and will transmit range safety commands
selected by either the RSO or the computer. Normal transmission
is resumed after transmission of the high-priority command.
To initiate a command from his console, the RSO would actuate
a hooded toggle switch. The output of the encoder is then routed
in parallel form to a tone remoting transmitter that processes the
message for transmission over the 5 mi distance to the transmitter
site. A tone-remoting receiver at the transmitter site demodulates
the message and feeds it to a modulator that converts the parallel
information to the high-alphabet, l l-character format. The I1dual-tone bursts are then fed to the AN/FRW-2A transmitter
system and to the vehicle. For reliability, a completely redundantbackup system is provided, with a continuously monitoring error
detector that provides automatic transfer to the backup system if
the primary chain should fail.
DOWNRANGE REMOTE SITES.
When the vehicle is below the radio horizon of Cape Kennedy,
the RF transmission will be made from one of the downrangesites. All downrange sites are connected by cable: when the RSO
presses a command switch on his console, the command pulse will
be transmitted over the cable via the supervisory control system.
(This system wilt be replaced by a digital remoting system in thenear future.) All sites will receive the command, and the transmitter
on the air at the moment the command is received will transmit
the message to the vehicle. (To keep the onboard receiver captured,one transmitter is always on the air.)
RANGE SAFETY SYSTEM MEASUREMENTS.
Ten measurements (figure 1-11) of the S-IB and S-IVB range safety
systems are taken during flight and telemetered to ground receivingstations where they are recorded for postflight evaluation. The four
command-receiver, low-level signal-strength measurements are
monitored in real time on the auxiliary pad safety panel at KSC.
The two S-IVB stage receiver strength signals are flight control
measurements that require real-time monitoring by Mission Control
Center at Houston, Texas. The S-IB stage telemetry system trans-
mits the S-IB stage range safety measurement data continuously
from liftoff until S-IB stage impact. The S-IVB stage telemetry
system transmits S-IVB range safety system data from liftoff
through orbital coast. All S-IVB signals should be zero-volt signals
after the range safety officer safes the S-IVB range safety system.The safe command removes power from the receiver and decoders
to deactivate the system after orbit has been attained and the rangesafety system is no longer necessary.
1-12
Section General Description
STAGE MEASNO.
S-IB K65-13
S-18 K66-13
S-IB VMS05.- 13 [_;_
S-IB VMS08-13 _>
S-IVB K98-411
S-IVB K99-411
S-IVB M30-411 [_
S-IVB M31,411
S-,VB V NE7-411_"_
S-IVB VN62-411 1_22_
TITLE
EVENT - Pv/S 1 CUTOFF/DEST IND
EVENT - R/S 2 CUTOFF/DEST IND
R/S RCVR ! LOW-LEVEL SIG STRENGTH
R/S RCVR 2 LOW-LEVEL SIG STRENGTH
EVENT - R/S ! ARM/CUTOFF DEST lIND
EVENT - R/S 2 ARM/CUTOFF DEST IND
VOLT, FU 1 EBW, RANGE SAFETY
VOLT, FU 2 EBW, RANGE SAFETY
R/S RCVE 1 LOW-LEVEL SIG-STRENGTH
R/S RCVR 2 LOW-LEVEL SIC--STRENGTH
MONITORED IN REAL TIME AT KSC
MONITORED IN REAL TIME AT MCC
Figure 1-11
RANGE I
OTOSV I
0TOSV I
OTOSV ]
0TOSV J
0TOSV I
0TO5V I
0TOSV I
OTO 5V
0TOSV
0TO5V
C-H 14486-1
1-13
SECTION II
TABLE OF CONTENTS
Introduction ........................................................................ 2-1
Mission Description ............................................................ 2-1
Propulsion Performance .................................................... 2-7
Separation Dynamics .......................................................... 2-12Mass Characteristics .......................................................... 2-16
Tracking Coverage ............................................................ 2-18
INTRODUCTION.
This section contains performance data that represent the SaturnIB manned missions for the Skylab program. The launch vehicle
data presented is for SA-206 or SA-207, but may be consideredtypical of the Saturn IB launch vehicles assigned to the Skylab
program• The flight sequence illustrated in Figure 2-1 is typicalof the launch vehicle mission profile• Figure 2-2 summarizes launch
vehicle performance characteristics of SA-206, the first Saturn IBvehicle scheduled for a Skylab flight•
MISSION DESCRIPTION.
The Skylab (SL) program consists of four missions designed to
achieve long duration space flights of men and systems and to
FLIGHT SEQUENCE OF EVENTS
FILGHT ROGRAMITIME
HR: MIN: (SEC) TIME
SEC) (SEQ
,0:0:17.0
• 0:0:03. t
0:0:00.0
0:0:00.2
0..0:10.2
0:0:58.9
0:1:13.6
0:1:40.C
0:2:00.G
0:2:10._
0:2:11.1
0:2=14._
0:2:17._
0:2:20._
0:2:21 .S
0:2:22.{
0:2:23.,
0:2:25.,
0:2:26,
0:2:29.,
0:2:33.
0:2:45.,
0:2:45
0:2:50
0:3:02
, 0:5:46"
i 0 t7:48
I 0:9:41
0:9:42
0:9:51
- 17.0_
- 3.1C
0.0C
0.2(
10.2(
58.87
73,61
100,0(
120.0(
130.5(
131.I:
134.6:
137.6:
140.&
141.95
142.03
143.35
145.70
146.75
149.35
153.95
165,24
165.65
170,65
182.65
346.75
468.75
581.93
582.13
591.93
EVENT
....... , GUIDANCE REFERENCE RELEASE (GRR).
....... jINITIATE S-IB MAINSTAGE IGNITION SEQUENCE.
....... FIRST MOTION.
(0.0) 1 I LIFT-OFF SIGNAL; INITIATE TIME BASE 1.
(10.0) 1 INITIATE PITCH AND ROLL MANEUVERS........ MACH ONE.
....... =MAXIMUM DYNAMtC PRESSURE.
( 99.8) 1 ICONTROL GAIN SWITCH POINT.
(119,8) 1 JCONTROL GAIN SWITCH POINT.
(130.3) 1 JTILT ARREST.
(130.9) I I
( 0.0) 21
( 3.0) 2 !
( 0.0) 3
( 1.3) 3
( 1.4)3
C 2.7)3
( 8.7)3
( 13.3)3
( 30.0)3
(42.0) 3
(206.1) 3
(328.1) 3
!.°..72:
ENABLE S-IB PROPELLANT LEVEL SENSORS.
LEVEL SENSOR ACTIVATION; INITIATE TIME BASE
INBOARD ENGINE CUTOFF (IECO).
OUTBOARD ENGINE CUTOFF iOECO).
INITIATE TIME BASE 3,
SEPARATION SIGNAL.
S-IB/S-IVB PHYSICAL SEPARATION;
CONTROL GAIN SWITCH POINT.
J-2 ENGINE START COMMAND,
ULLAGE BURN OUT,
90% J-2 THRUST LEVEL.
COMMAND 5.5:1 EMR.
JETTISON ULLAGE ROCKET MOTORS.
DYNAMIC PRESSURE = 1 PSF.
LES JETTISON.COMMAND ACTIVE GUIDANCE INITIATION.
CONTROL GAIN SWITCH POINT.
CONTROL GAIN SWITCH POINT.
COMMAND EMR SHIFT TO 4.8:1.
GUIDANCE CUTOFF SIGNAL (GCS).
INITIATE TIME BASE 4.
RBIT INSERTION, CH.2006_
Figure 2-1
S-IB STAGE
AVERAGE LONGITUDINAL SEA LEVEL THRUST (LBF)
H-1 ENGINE TURBINE TOTAL
ENGINE !1 207575 663 208238
ENGINE #2 206923 669 207592
ENGINE #3 207031 658 207689
ENGINE #4 207234 706 207940
E NGNE #5 208745 665 209410
ENG INE t6 207955 665 208620
ENGINE #7 207968 667 208635
ENGINE 18 207680 656 208336
TOTAL AVERAGE LONGITUDINAL SEA LEVEL
THRUST 1666460
FLIGHT TIME INTERVAL: 0.0 - 137.648 SEC (IECO)
_V = [77T. @ it=()) - WT @ (t = 137.648) - *WAux] / 137.648
= 6346.72 (LBM/SEC) *WAux: FROST 1100 LBMSEAL PURGE 6 LBM
ISP = F/_
= 262.S7 ($EC)
S-IVB STAGE
HIGH THRUST LEVEL FLIGHT TIME INTERVAL:
LOW THRUST LEVEL FUGHT TIME INTERVAL:
FUEL ADDITIVE 27 LBM
TOTAL
15t.00- 469.30SEC
469.50- 581.93 SEC
AVERAGE VALUES
HIGH THRUST LOW THRUST
LEVEL LEVEL
VACUUM THRUST (LBF) 229,714. 198,047.
FLOWRATE (LBM,/S EC) 543.87 465.69
SPECIFIC IMPULSE (SEC) 422.37 425.28
CH_
Figure 2-2
perform scientific investigations in earth orbit. The first mission,
SL-I, will utilize a Saturn V launch vehicle to place the Saturn
Workshop (SWS) into a 233.8 NM orbit inclined at 50 degrees•
Currently, four Saturn IB launch vehicles are assigned to the Skylab
program• SA-206 is assigned for the SL-2 mission, SA-207 the SL-3mission, and SA-208 the SL-4 mission. SA-209 is assigned to the
program as a backup launch vehicle. The SL-2, SL-3, and SL-4missions are scheduled for launch at 90 day intervals, transporting
three man crews to the SWS for operations of 28 to 56 days duration•
Mission requirement for the Saturn IB launch vehicle is insertionof the Command and Service Modules (CSM) in an 81 by 120
NM elliptical orbit which is co-planar with the SWS orbit. After
separation, the CSM propulsion systems will be utilized to achieveorbit transfer, rendezvous and docking with the SWS.
TRAJECTORY.
The trajectory data presented were extracted from the latest pub-
lished reports for the SA-206 and SA-207 launch vehicles. It covers
all phases of flight for which the launch vehicle has a functional
requirement. These tabulations and curves reflect three flight
phases. The S-IB stage flight phase starts at guidance referencerelease, which is assumed to occur 17 sec before first motion. The
S-IB stage flight phase ends at physical separation of the S-IB
2-1
E
Section II Performance
EVENT
GUIDANCE REF. RELEASE
FIRST MOTION
MACH ONE
MAX. DYN, PRESSURE
TILT ARREST
INBOARD ENGINE CUTOFF
OUTBOARD ENGINE CUTOFF
S-I8/5-1VB PHYSICAL SEP.
J-2 ENG. START COMMAND
ULLAGE CASE JETTISON
LES JETTISON
IGM INITIATION
EMR SHIFT, 5.5:1 TO 4.8:1
GUIDANCE C/O SIGNAL
ORBIT INSERTION
FLIGHT
TIME
(HR:MIN:SEC)
-00:00..05.0
00:00:00.0
00:00:58.9
0o:0h13.6
00..02:10.5
00..02:17.6
00..02:20.6
00=02:22.0
00:02:23.4
00:02:34.0
00:02:45.6
00:02:51.0
00:07:49.5
00:09:41.9
00..09:51.9
EARTH FIXED
DECLINATION LONGITUDE
(DEG) (DEG) W.
28.466 80.621
28.466 80.62t
28.478 80.607
28.497 80.586
28.767 80.272
28.839 80.189
28.872 80.150
28.887 80.132
28.902 80.114
29.021 79.974
29.157 79.813
29.221 79.738
34.950 72.168
38.727 66.106
39.115 65.418
ALTITUDE
(FT)
295.
295.
24,282.
40,760.
158,431.
179,884.
189,284.
193,605.
196,616.
228,898.
260,970.
274,862.
540,858.
519,714.
520,175.
VELOCITY
(FT,/SEC)
1,340.44
1,340.44
1,942.46
2,416.83
6,547.02
7,394.01
7,586.40
7,591.20
7,575.35
7,621.65
7,773.19
7,852.68
17,955.50
25,681.12
25,705.77
SPACE FIXED
FLIGHT PATHANGLE (DEG)
90.000
90.000
60.479
57.840
63.884
65.151
65.659
65.923
66.186
68.197
70.240
71.134
90.671
90.009
89.998
AZIMUTH
(DEG)
90.000
90.000
79.091
72.693
54.910
53.816
53.603
53.593
53.603
53.519
53.348
53.269
52.953
38.727
39. 303
Cfl.20070
Figure 2-3
stage from the S-IVB stage, assumed to occur approximately 1.4
sec after outboard engine cutoff signal. The S-IVB stage powered
flight phase extends from S-IB/S-IVB separation to orbit insertion.
Orbit insertion is defined as 10 sec after the guidance cutoff signal.
Guidance cutoff occurs when the space-fixed velocity equals a
prespecified value that, accounting for the additional velocity im-
parted by the J-2 engine thrust decay, provides a velocity of 25,705.8
ft/sec at orbit insertion. The S-IVB/IU orbital flight phase is de-
fined as from orbit insertion to the predicted loss of attitude control
capability at approximately 7V2 hours of flight. See figures 2-3
through 2-15 for detailed trajectory information.
Trajectory Dispersions.
Since performance predictions and associated calculations are sub-
ject to certain tolerances, a dispersion analysis is conducted toestablish realistic deviation limits. The error sources considered for
the analysis are those associated with predictions of vehicle charac-
teristics, vehicle systems performance, and flight environment. Fig-
ure 2-16 summarizes the trajectory dispersion envelopes at S-IB/S-
IVB stage separation and orbit insertion. These values reflect the
combined S-IB and S-IVB stage three-sigma deviations using the
root-sum-square (RSS) technique as follows:
+RSS = _/Z (+A P)-'
-RSS = _f_(--AP)'
where
A P = Perturbed Parameter - Nominal Parameter
The RSS technique is also utilized to determine the Flight Perfor-
mance Reserve (FPR) propellant required to ensure a three-sigma
probability of achieving the desired S-IVB end conditions of flight.
SL-2 ALTITUDE VS GROUND RANGE
2-2
/
/f
200
j _ S-IB STAGERE-ENTRY
4OO 6OO
GROUND RANGE (NM)
80O
Figure 2-4
10O0
CH-2Cg71
Section II Performance
SL-2FLIGHTPATHANGLEHISTORIES
500
400
300
200
100
//
/
//,/
/f
//
G.J0 1:40 3:20 5:00
\
\/-- - S-18 STAGE
/ RE-tNTRY --
\6:40 8:20 10..00
FLIGHT TIME (MIN:SEC)
CH-20072
Figure 2-5
4=
//
1:40
I j I /I
•"O- "O" SPACE FIXED /-- EARTH FIXED
t
S-IB RE-ENTRY
f
3:20 5:00 6:40
FLIGHT TIME (MIN:SEC)
8:20
Figure 2-7
10:00
SL-2 VELOCITY IISTORIES
24000,
20000
016000
12000,
> 80OO
400O
O
0
"_ --O- SPACE FIXED-- EARTH FIXED
d
1:40 3:20 5:00 6:40
FLIGHT TIME (MIN:SEC)
;r
IS-IB STAGE --
RE-ENTRY
8:20 10:0C
CH.20073
Figure 2-6
120.
80.
i/
0
i
/
//!
II
EMR SHIFT 7
"V/
/
1:40 3:20 5:00 6:40
FLIGHT TIME (MIN:SEC)
8:20
Figure 2-8
/
10:00
CH-20075
2-3
Section II Performance
;L-2 rAW ANGLE OF ATTACK HISTORY
800.
600.
_u
Z
200.
J0 0:40
\1:20 2:00
FLIGHT TIME (MIN:SEC)
Figure 2-9
2:40 3:20
CH-20076
1:20
).80
).40 I_
0
).40
).80
.20
.600 1:40
W\^\/\ ^
3:20 5:00 6:40
FLIGHT TIME (MIN:SEC)
8:20
Figure 2-11
CH-2_B
St-2 PITCH ANGLE OF ATTACK HISTORY SL-2 TOTAL ANGLE OF ATTACK HISTORY
20.
16.
0
12.
8.
0
z 4.
0
-4.
1:40
f ,
.I I
J
/
I3:20 5:00 6:40
FLIGHT TIME (MIN:SEC)
\
\\
8:20 10.-00
CH-20077
Fiqure 2-10
4.
A_,_I_
00 1:40
_l
/
3:20 5:00 6:40
FLIGHT TtME (MIN:SEC)
\\\
V
r8:20 10:00
CH-20079
Figure 2-12
2-4
SectionIIPerformance
FLIGHT TIME (t) : OECO + 1.379 SECONDS
RADIUS (R):ALTITUDE:
SPACE FIXED VELOCITY (V):
SPACE FIXED PATH ANGLE (0):
SPACE FIXED FLIGHT AZIMUTH ( AZI):
EARTH FIXED FLIGHT AZIMUTH (AZE):
GEOCENTRIC DECLINATION ( 6):
GEODETIC LATITUDE (¢):
LONGITUDE (),)= (WEST)
142.027 (SEC)
21,102,868. (FT)
193,605. (FT)
7,591.19 (FT/S)65.923 (DEG)
53.593 (DEG)
45.804 (DEG)
28.887 (DEG)
29.050 (DEG)
80.I32 (DEG)
SPACE FIXED POSITION AND VELOCITY COMPONENTS
XS 21,099,643. (FT)
Ys 178,280. (FT)
Zs 322,913. (FT)
)_s 2,983.85 (FT/S)
_'s 901.80 (FT/'S)Zs 6,921.67 (FT/_)
VEHICLE ATTITUDE AND ATTITUDE RATE
PITCH ATTITUDE ANGLE (¢p)= -63.251 (DEG)
YAW ATTITUDE ANGLE (¢y): - 0.073 (DEG)
ROLL ATTITUDE ANGLE (¢ r)t 0.007 (DEG)• 0.026 (DEG/S)PITCH RATE (¢p):
YAW RATE (¢y): 0.009 (DEG/S)
ROLL RATE (¢ r): 0.007 (DEG/S)
CH-20080
Figure 2-13
SL-2 S-IVB STAGE END CONDITIONS OF FLIGHT
AT GUIDANCE CUTOFF SIGNAL'
FLIGHT TIME (t): GCS 581.926 (SEC)
RADIUS (R): 21,417,871. (FT)
ALTITUDE: 519,714. (FT)
SPACE FIXED VELOCITY (V): 25,681.12 (FT/S)
SPACE FIXED FLIGHT PATH ANGLE (®): 90.009 (DEG)
SPACE RiXED FLIGHT AZIMUTH (AZI): 55.485 (DEG)
EARTH FIXED FLIGHT AZIMUTH (AZE): 53.882 (DEG)
GEOCENTRIC DECLINATION ( 6 ): 38.727 (DEG)
GEODETIC LATITUDE (¢): 38.915 (DEG)
LONGITUDE (X): (WEST) 66.106 (DEG)
INCLINATION (i): 49.999 (DEG)
DESCENDING NODE ARGUMENT (®): 154.672 (DEG)
INERTIAL RANGE ANGLE: 17.416 (DEG)
SPACE FIXED POSITION AND VELOCITY COMPONENTS
Xs 20,447,603. (FT)
Ys 456,794. (FT)
Z$ 6,357,044, (FT)
Xs -7,628.17 (FT/S)
Ys 2.79 (FT/S)
Zs 24,522.06 (FT/'S)
VEHICLE ATTITUDE ANGLES
PITCH ATTITUDE ANGLE (¢p) -99.994 (DEG)
YAW ATTITUDE ANGLE (¢y) -2,899 (DEG)ROLL ATTITUDE ANGLE (¢ r) 0.258 (DEG)
OSCULATING CONIC PARAMETERS
* PERIGEE ALTITUDE 80.97 (NM)
* APOGEE ALTITUDE 105.69 (NM)
ECCENTRICITY 0.0035
SEMI-MAJOR AXIS 3,537.26 (NM)
TRUE ANOMALY 357.33 (DEG)
PERIOD 87.95 (MIN)
*REFERENCED TO EQUATORIAL RADIUS (3,443.93 NM)
CH.20061
Figure 2-14
Engine-Out Capability.
At any time during the S-IB stage boost powered flight after 3.0
sec, there exists the possibility of having one of the eight H- l engines
shut down prematurely. The mission profile must then be completed
FLIGHT TIME (t): ORBIT INSERTION 591.926 (SEC)
RADIUS (R): 21,417,867. (FT)ALTITUDE: 520, t76. (FT)
SPACE FIXED VELOCITY (V): 25,705,56 (FT/S)SPACE FIXED FLIGHT PATH ANGLE (e): 89.998 (DEG)
SPACE FIXED FLIGHT AZIMUTH (AZI): 55.941 (DEG)
EARTH FIXED FLIGHT AZIMUTH (AZE): 54.367 . (DEG)
GEOCENTRIC DECLINATION (6): 39.115 (DEG)
GEODETIC LATITUDE (¢): 39.303 (DEG)
LONGITUDE (x): (WEST) 65.418 (DEG)
INCLINATION (;): 50.000 (DEG)
DESCENDING NODE ARGUMENT (0): 154.675 (DEG)
INERTIAL RANGE ANGLE: 18.103 (DEG)
SPACE FIXED POSITION AND VELOCITY COMPONENTS
Xs 20,369,815. (FT)
Ys 456,777. (FT)
Zs 6,602,044. (FT)
)_s -7,924.82 (PT/S)
_'s - 4.97 (FT/_)
_[s 24,453.69 (FT/S)
VEHICLE ATTITUDE ANGLES
PITCH ATTITUDE ANGLE (¢p) -99.996 (DEG)
YAW ATTITUDE ANGLE (0y) -2.906 (DEG)ROLL ATTITUDE ANGLE (¢r) .102 (DEG)
OSCULATING CONIC PARAMETERS
*PERIGEE ALTITUDE 80.99 (NM)
*APOGEE ALTITUDE 119.37 (NM)
ECCENTRICITY 0.0054
SEMI-MAJOR AXIS 3,544.11 (NM)
TRUE ANOMALY 0.30 (DEG)
PERIOD 88.20 (MIN)
*REFERENCED TO EQUATORIAL RADIUS (3,443.93NM)
CH.200_
Figure 2-15
with the remaining seven engines. The S-IB stage steering com-
mands are the body attitude Euler angles preprogrammed in the
Launch Vehicle Digital Computer (LVDC) as functions of time
only (i.e., the steering program is open-loop). The pitch attitude
steering program (CHI) is designed to enforce a near-zero angle
of attack time history throughout the high dynamic pressure regime
of normal vehicle flight. At a preprogrammed time of approximately
10 seconds prior to outboard engine cutoff (OECO), the pitch
steering command is arrested and remains constant for the re-
mainder of the S-IB stage flight and for the first 29 sec of S-IVB
stage flight.
Chi-Freeze Steering.
Subsequent to an engine failure, the vehicle guidance is switched
to the chi-freeze steering mode. In the chi-freeze steering mode,the value of the commanded pitch attitude is frozen, upon engine
failure, for an incremental duration and then the nominal program
(displaced in time) is resumed until chi-arrest. The chi-freeze inter-
val is a function of engine failure time as depicted in figure 2-17.
Because extended periods of vertical or near-vertical flight are
objectionable near the launch complex, the chi-freeze mode is
inhibited during the first 30 sec of flight. For an engine failure
during the inhibited period, the pitch attitude follows the nominal
program until 30 sec of flight at which time the proper (i.e., as
a function of engine out time) chi-freeze interval is initiated. For
engine failures prior to 110 sec, chi-arrest is initiated as the nominal
time plus the chi-freeze interval minus 20 sec. The deviation in
the pitch attitude command history due to utilization of the chi-
freeze steering mode for an H-1 engine failure at liftoff is illustrated
in figure 2-18.
2-5
Section II Performance
THREE-SIGMA TRAJECTORYDISPERSION ENVELOPES(TYPICAL
TRAJECTORY PARAMETER
FLIGHT TIME (SEC)
ALTITUDE (FT)
SPACE FIXED VELOCITY (FT/'SEC)
SPACE-FIXED FLIGHT PATH ANGLE (DEG)
GROUND RANGE (FT)
SPACE-FIXED POSITION VECTOR (FT) IY
X
SPACE-FIXED VELOCITY VECTOR (FT/SEC)
_ITCH
VEHICLE ATTITUDE (DEG) L_OL_YALW
THREE-SIGMA DISPERSION ENVELOPES
S-IB/1VB SEPARATION ORBIT INSERTION
+RSS -RSS +RSS -RSS
4.26
8,494.
133.60
2.281
17,972.
8,602.
12,602.
21,962.
233.53
224.93
194.09
1.538
1.522
3.648
2.94
7,894.
144.16
2.008
14,088.
8,150.
9,30.
16,434.
279.40
212.01201.44
1.534
1.521
3.648
I1.75
1,722.
5.22
0.018
129,485.
44,347.
16,749.
134,708.
159.28
10.50
53.97
2.527
2.515
10.67
1,749.
5.45
0.018
126,663.
46,266.
17,231.
131,250.
163.09
10.50
56.27
2.481
2.280
NOT APPLICABLE DUE TO APS CONTROL LIMIT OF APPROXIMATELY ONE DEGREE ERROR.
CH.20083
Figure 2-16
PITCH ATTITUDE COMMAND HISTORY COMP_,RISONCHI-FREEZE INTERVAL
2O
15
Z
o
_1o
5
00 20 40 60
TIME OF ENGINE FAILURE (SEC)
CH-2_)84
Figure 2-17
Performance Capability.
For a planar flight launch and nominal performance, the predicted
weight-in-orbit capability for the SA-206 vehicle is 67,790 Ibm.
This provides approximately 3,400 lbm. of excess useable S-IVB
2-6
2
8
0
-20
-40
-80
NOMI NAL
ENGINE OUT
AT 0 SEC
40 80 120
FLIGHT TIME (SEC)
160
CH-20085
Figure 2-18
stage propellant. This propellant may be utilized to compensatefor low performance, perform S-IVB stage yaw steering to provide
launch windows, or improve engine-out capability. Propellant re-quired to compensate for three-sigma low performance (FPR) is
2,200 Ibm. Propellant requirements to compensate for single H-lengine failures are provided in figure 2-19. The launch window
performance requirements are discussed in Section IX.
PROPULSION PERFORMANCE. S-IB STAGE PROPELLENT DEPLETION LEVELS (TYPICAL)
The predicted launch vehicle propulsion performance, as presented,
is derived from computer simulations of the S-IB and S-IVB pro-
pulsiOn system performance using data from stage static tests, singleengine acceptance tests, and previous flight tests.
S-IB STAGE.
Performance data of the S-IB stage is extracted from prediction
data for the SL-2 mission. Stipulated criteria of the prediction are:
1) Launch from LC-39B of the AFETR;
2) Terminal delivery into an 81 x 120 NM phasing orbit.
Engine Simulation.
The predicted performance for the eight, 205K thrust H-I engines
was calculated from ground test data using a table of influence
coefficients (partials) to estimate the performance with the pump
inlet pressures and propellant densities expected during flight.
Basically, the engines are characterized by Rocketdyne single engine
acceptance test data modified by empirical factors derived from
previous flight data. Since S-IB-6 is the first stage to use 205K
thrust engines and no previous flight history exists, the magnitudes
of the factors were influenced by the results of the stage statictests conducted at MSFC. The factors are somewhat smaller than
would be used for 200K engine stages.
H-1 ENGINE FAILURE PERFORMANCE EFFECTS
6
r-_
Q)_ 3
_z0"_. 2
===.1
00
---,.....
10 20 30 40 50 60 70 80 90 100
TIME OF ENGINE FAILURE (SEC)
Cfl-20086
Figure 2-19
Section II Performance
EXPECTED LEVEL
AT END OF
THRUST DECAY
/ .ox I LEVEL I LOX I I\ TAN; I/ LEVItt
27._ IN. IN.
L.-.- _ J BIASING -""'--_ _ FUEL DEPLETION PROBE
J FUEL _ MINIMUM LEVEL AT OUT-
_ _ ROA,DENG,NECUTOFFEXPECTEDLEVEL--'-'--...._._ _ _i1_AT END OF _ _ k_l _"-- MI NI MUM LEVEL
T._STOECAV_2_.__H _ _ '_ AT#_%AY
OUTBOARD INBOARD OUTBOARD
ENGINE ENGINE ENGINE
_ EXPECTED LEVEL IS ASSUMEDTO BE RESULTS OFLOX STARVATION CUTOFF. MINIMUM FUEL
LEVELS RESULT WITH FUEL _EPLETION CUTOFFS CH-20_7
Figure 2-20
Pressurization and Propellant Data.
The predicted lox tank and fuel tank pressurization historiesexpected for flight were calculated from data obtained from the
MSFC stage static tests.
The predicted fuel density was determined from temper-
ature-density chemical analyses of the fuel that will be used and
the expected fuel temperature at launch. The fuel temperature isbased on the ambient temperature and an appropriate chilldown
due to lox tank proximity expected for the time of year of launch.
The lox density was determined from the mean bulk temperature
calculated from the wind velocities, ambient temperature, ambient
pressure, and humidity expected at the time of launch.
Propellant Utilization.
The predicted amount of lox loaded is established as a full load
to the minimum ullage of 1.5 percent. The predicted fuel load
is calculated as the amount required to deplete all of the usable
lox load with the performance expected during flight.
The predicted fuel load is not necessarily the amount of fuel loaded
on the stage at the time of launch. Propellant load tables are
furnished to KSC to provide fuel load data for the actual fuel
temperature at launch time. No adjustments to the predicted lox
loads are attempted.
The amount of lox trapped residual (approximately 2600 Ibm)is determined as the lox in the suction lines to the main valves
of the inboard engine, a few gallons trapped in the center lox
tank sump, and approximately 70 gal in the outboard engines.
The amount of trapped fuel residual (3070 Ibm) is determinedas the amount of fuel remaining if the engines are cutoff by the
fuel depletion sensors located in the sumps of F-2 and F-4 (figure2-20).
Included in the fuel load is a 1550 Ibm of fuel bias which provides
a nominal fuel residual of 4620 Ibm. The fuel bias is provided
to minimize the amount of propellant residuals caused by deviations
from predicted consumption ratios and loading inaccuracies.
2-7
Section II Performance
S-IB Stage Predicted Performance.
Figure 2-21 shows the ambient pressure profile during powered
flight. Individual S-IB stage propulsion system performance para-
meter variations are shown in figures 2-22 through 2-31.
S-IBtOX ULLAGEPRESSURE
AMBIENTPRESSUREHISTORY
12
8
\\
\
\\
\\
\\
0 0:20 0:40 1:00 1:20 1:40 2:00 2:20 2:40 3t00
FLIGHT TIME (MIN:SEC)
CH-20_8
Figure 2-21
70
6O
4O
0 0:20 0:40 hoo I;20 1:40 2:00 2:20 2:40 3:0C
FLIGHT TIME (MIN:SEC)
CH.20090
Figure 2-23
S-IBLOXFLOWRATES-IBFUELULLAGEPRESSURE
2-8
0 0_20 0:40 I :OO 1:20 1:40 2:00 2=20 2:40 3:00
FLIGHT TIME (MIN:SEC)
CH-2GO_I9
Figure 2-22
5200
48OO
44OO
4OOO
32OO
280O
24OO
2OOO
18OOq 0:20 0:40 1.-00 1:20 1,40 2:00 2=20 2:40 3:OO
FLIGHT TIME (MIN:SEC)
CH-20091
Ftgure 2-24
Section II Performance
S-IB FUEL FLOWRATE
'600
1400
_200
;000
8OO
6OO
400
2OO
000
8OO
600
4OO
200
00 0:20 0:40 1:00 1:20 1:40 2:00 2=20 2:40 3:00
FLIGHT TIME (MIN:SEC)
CH-20092
Figure 2-25
36oo!!
3200 I
2800 I
_. 24OO1
20001
==
> 12001
I I
800 :
I I400,
0 0=20 o:4o :._' i,20 1:40 2:oo 2=20 2:40 s=oo
FhGHT TIME (MIN:SEC)
CH.20094
Figure 2-27
klB SPECIFIC IMPULSES-IB TOTAL PROPELLANT FLOWRATE
8
6
4
2
0
8
6 -
4
2
0 0 0:20 0:40 1:00 1:20 1:40 2:00 2:20 2:40 3:00
FLIGHT TIME (MIN:SEC)
CH.20093
Figure 2-26
38O
36O
34O
320
0oo
280
260
24O
I"
220
200,0 0:20 0:40 h00 1:20 h40 2:00 2:20 2:40 3:00
FLIGHT TIME (MIN:SEC)
CH._095
Figure 2-28
2-9
SectionIIPerformance
S-IBTHRUSTDECAYS-IB MIXTURE RATIO
2.4
2.3
Z.2
2.1
2.00
_j
0:20 0:40 1:00 1:20 1:40 2:00 2:20
FLIGHT TIME (MIN:SEC)
Figure 2-29
2:40 3:00
CH.20_96
r2oo_
IO60 4 OUTBOARD ENGINES
2:16 2:17 2:18 2:19 2:20 2:21 2:22 2:23 2:24 2:2S
FLIGHT TIME (MIN:SEC)
CH.20098
Figure 2-31
S-IB THRUST BUILDUP
80C
60C
40C
20C
8O6
6O6
40C
2O6
2-1 0
C_-3.S -3.0
/J
-2.5 -2.0 -1.5 -1.6 -0.5
FLIGHT TiME (SEC)
0.0 0.5 1.0
CH-20097
Figure 2-30
S-IVB STAGE.
The S-IVB stage data were extracted from reference propulsion
flight performance predictions for a 81 by 120 NM orbital mission.
This prediction is based on the following stage propellant utiliza-tion:
a. Total Iox load of 195,972 Ibm.
b. Total fuel load of 37,900 Ibm.
c. Total lox consumed by engine and boiloff during flight of 192,621Ibm.
d. Total fuel consumed by engine and for tank pressurization
during flight of 36,016 Ibm.
e. Lox residual of 3,351 Ibm.
f. Fuel residual of 1,884 Ibm.
Figures 2-32 through 2-38 show stage propulsion system perfor-
mance parameter variations.
Section II Performance
J-2 ENGINE THRUST BUILDUP
220
200
180
160
140
120
100
80
60
4O
20
f--
/NOTES:
I. MIXTURE RATIO
CONTROL VALVE
IN 4.8 EMR POSITION
2. ENGINE START
COMMAND AT 2:23.3
/
2:23 2:24 2:25 2:26 2:27 2:28 2:29 2:30
VEHICLE FLIGHT TIME (MIN:SEC)
Figure 2-32
220
2OO
180
160
140
120
100
80
6O
4O
20
9:41
NOTES:
I. MIXTURE RATIO
CONTROL VALVE
IN 4.8 EMR POSITION
2. GUIDANCE CUTOFF --
SIGNAL AT 9:41.9
\9:42 9:43
VEHICLE FLIGHT TIME (MIN:SEC)
Figure 2-34
9:44
CH.20101
I-IVB SPECIFIC IMPULSES-IVB .ONGITUDINAL THRUST
26O
24O
220
200
180
160
1:40
11
/---ENGINE START
COMMAND
T=2:23.3
GCS JT=9:41.9 -
3:20 5:00 6:40 8:20
PREDICTED TIME FROM VEHICLE FIRST MOTION (MIN:SEC)
Figure 2-33
10:0C
CH-2010O
434
432
43(3
42_
426
424
42;
42(
41E1:40 3:20 5:00 6:40 8:20 10:00
TIME FROM VEHICLE FIRST MOTION (MIN:SEC)
Figure 2-33
CH-20102
2-11
SectionII Performance
S-IVB ENGINEMIXTURERATIOS-IVB LOX FLOWRATE
Figure 2-36
5.8
5.6
5.4
5.2
5.0
4.8
4.6
4'4 L4.2
1:40 3:20 5:00 6:40 8:20 10:00
TIME FROM VEHICLE FIRST MOTION (MIN:SEC)
CH-20105
Piqure 2-38
;-IVB FUELFLOWRATE
84G
82
8o
_ 78
76
2-12
=_' ENGtNE STARTCOMMAND
T=2:23.372 i I
1:40 3=20 5:00
| I
I I
' I
6:40 8:20
TIME FROM VEHICLE FIRST MOTION (MIN:SEC)
10:00
Figure 2-37
SEPARATION DYNAMICS.
The latest launch vehicle dynamics analyses indicate that no prob-
lem exists for S-IB/S-IVB stage separation motion. Potential prob-lems considered ate relative lateral motion of the J-2 bell and S-IVB
aft interstage wall during physical separation, and S-[VB post
separation controllability. The probability (cumulative distribution)of the J-2 beil clearing the S-IVB aft interstage wall in the case
of a single retro failure in combination with stage separation tolera-
nces is in excess of 3-sigma.
Both potential separation problems of J-2 bell interstage collision
and S-IVB stage controllability are mainly affected (assuming no
retro failures) by large _erodynamic moments or attitude rates
existing at first stage boost flight termination. These two problems
are minimized by first stage boost trajectory shaping. A nose-over
maneuver initiated approximately 95 sec into flight and a tilt arrest
initiated approximately 131 sec into flight results in acceptable
levels of dynamic pressure, a small angle-of-attack, and attituderates that are essentially zero at S-IB/S-IVB separation.
The main contributor to the physical separation of the S-IB stage
from the S-IVB stage is the thrust of the four retro rockets. To
o
x
z
1000
S00
6OO
4OO
2OO
0
160
120
80
4O
0
J 1 f,,l
r"J_ -- RE_RONTH_US/
I I I ITHRUST DECAY (4 ENG)
I I/- "-'r- s-,BRETROOUTIY --N AL,'LUMEI
/., %./_1..,..... _.ALPiUMErULLAGE THRUST (3 ENG)
ULLAGE IGNITION
SEPARATION SIGNAL
RETRO IGNITIONJ-2 START COMMAND J-2 90% THRUST
I .... I , ,
1 2 3 4 5 6 7 8
TiME FROM OECO (SEC)
CH.2010_
Figure 2-39
a very slight degree, the three ullage rockets also contribute to
the physical separation. Proper phasing of the retro thrust with
respect to the separation signal and H-1 thrust decay is necessary
for successful staging and is shown in Figure 2-39. Impingement
of the retro rocket plumes on the vehicle creates pressure distribu-tions on the surface of the S-IB/S-IVB interstage and lower S-IVB
stage. If a retro rocket fails to ignite, these pressure distributions
then become asymmetric thereby causing imbalanced forces to act
on the stages as shown in Figures 2-40 and 2-41. This imbalanced
force condition constitutes a potential S-IB/S-IVB collision hazard.
Figure 2-39 indicates that the S-IVB stage is without effective J-2
control thrust for approximately 4.5 sec after physical separation
from the S-IB stage. It is during this time interval that c WB stage
dynamic transients can become excessively large.
The S-IB/S-IVB relative motion resulting from each of four retro
rocket failures in combination with stage separation tolerances,
subsequent to a nominal S-IB boost flight, is analyzed in order
Section II Performance
to ascertain successful retro-out staging probability. The quoted
probabilities are defined by the probability law:
4
P = Z P_ Pi*
i=]
where: P - probability of successful separation with one retrorocket failed
P_ -- probability that retro rocket number "i" is the onewhich failed
E*- probability of successful separation with retro rocketnumber "i" failed
The P_* probabilities quoted pertain to the cumulative distribution
function. Each P_* is determined by root-sum-squaring the incre-mental lateral travel due to each tolerance with retro rocket number
"i" failed. Those stage separation tolerances which have the greatestinfluence on S-IB/S-IVB relative lateral motion are those which
create significant moments on the S-IB stage. Aerodynamic mo-
ments resulting from aerodynamic tolerances are not large enough
on either stage to be significant contributors to a potential S-IB/S-
IVB collision. The stage separation tolerances considered in the
retro-out collision analysis are, therefore, retro rocket thrust vari-
ation ('+-13.28%), retro rocket thrust misalignment ('+0.50°), andS-IB lateral CG deviation ('+ 1.1 in.).
The single retro rocket failure results are presented in figures 2-42
and 2-43. Figure 2-42 gives the lateral clearance of the undeflectedJ-2 bell bottom with the S-IVB aft interstage (at interstage exit
plane) for each of the four single retro rocket failures possible.These results are based upon all retro failures being simulated
during an otherwise nominal separation subsequent to a nominal
S-IB boost flight. The smallest lateral clearance is 0.434 meters
which results when retro no. 3 fails. Figure 2-43 presents the J-2
bell lateral drifts in profile view for nominal and retro-out condi-
tions. In addition 3 o-off-nominal drifts are presented for nominaland the worst case retro rocket failure condition (retro no. 3 out).
As depicted in the figure 2-43 the probability (cumulative distribu-tion) of the J-2 bell clearing the interstage for a single retro failure
in combination with stage separation tolerances is in excess of 3or.
The qa product (dynamic pressure times total angle-of-attack) at
physical separation may be used as an indicator of S-IVB stagecontrollability after separation. Large qa products result in large
dynamic responses in the S-IVB subsequent to physical separation.
These post-separation transients may cause the J-2 engine to har-
dover against the 7-degree gimbal limit. However, vehicle responseis such that the J-2 engine hardover exists for only a few seconds
(in the worst case) and S-IVB control is maintained throughout
flight.
The maximum S-IVB stage attitude and rate errors due to terminal
boost flight conditions at separation are to be found in pitch. Figure
2-44 illustrates the peak pitch attitude errors after physical separat-
ion for the S-IB stage engine failures analyzed.
2-13
Section II Performance
S-IB SEPERATIONS-IB MOMENT SCHEMATIC S-IB SEPARATIONS-IVB MOMENT SCHEMATIC
STA 11
CENTERLINE OF RETR_
THRUST CANTED 9.5 °
ENGINE - 3
STA 495.787 --
STA 377.226
STA
CENTERLINE OF H-1
THRUST CANTED 6 °
ENGI NE - 3
INTERSTAGE
POINT OF THRUST APPLICATION
OF 23400 N NORMAL FORCE DUE
TO PLUME IMPINGEMENT
.INE OF RE'fRO
THRUST CANTED 9.5 °,ENGINE- 1
H-1 NO. 3 H-1NO. 1
PLAN VIEW LOOKING AFT
PLANE
CENTERLINE OFH-1
THRUST CANTED6 o,
ENGINE-1CH-_t07
Figure 2-40
1STA 1970.251
ULLAGE NO. 1
YM
I 'PLAN VIEW LOOKING AFT
/STA 1406.737-- d -CG MAX NORMAL FORCE =980 N
• MAX AXIAL FORCE = 33000 N
/-- I-,-9_N
ULLAGE NO. 2 -_/' / /
POINT OF THRUST APPHCATIONOUETOPLUME,MR,NGEME STA.O. 0-- ySTA 1086.1_7 _1 --GIMBAL STATION PLANE
zi
Figure 2-41
One of the most important of the post physical separation responsesis the peak pitch attitude rate. As indicated in figure 2-.45, the
envelope of the peaks of these pitch attitude rates was 4.36 deg/sec
for engine no. 5 failure at liftoff. It can be seen that the 10 deg/sec
EDS abort limit (see Section III) for post separation controllability
is not violated by any of these acceptable malfunction modes.
J-2 engine gimbal deflections and time against the 7-degree limit
stop are shown in figure 2-46. The upper graph shows, for example.that failure of engine no. 5 at liftoff causes the J-2 to be on the
7-degree gimbal limit for 2.4 sec, while with a failure of that engine
subsequent to 4.5 sec, the engine does not reach this limit. The
lower graph illustrates the maximum deflections in pitch for allfailure times. It can be seen that engine nol 5 never reaches the
7-degree limit for a failure subsequent to 4.5 sec. Both graphs showthat an engine no. 2 failure will never cause the J-2 to reach the
gimbal limits.
RETRO FAILED LATERAL CLEARANCE
(METERS)
NO.I .471
NO.2 .469
NO.3 .434
NO.4 .435
[_:> LATERAL CLEARANCE OF THE UNDEFLECTED J-2 BELL BOTTOM
WITH THE S-IVB AFT INTERSTAGE AT THE INTERSTAGE EXIT PLANE.
Figure 2-42
CH-201091
2-14
Section II Performance
S-IVB PEAK PITCH ATTITUDE RATE FOR
S-IBENGINE1, 2, & 5 FAILURES
3
_2
0
(30) RETRO OUT.._
.251SEC _
1 2
I IS-IVB AFT INTERSTAGE WALL
IIIIIIIIIIIIIII IIIIIIIIIIIIIII illllllllllllll IIIIIIIIIIIIIII lllmlllllllll IIIIIII
_ "_ RErRO bUT7 ' I(3o, NAL
......... .94 SEC
4 5
FORWARD TRANSLATION (M)
Figure 2-43
CH-2011{
11
t-_- 6
4 _
_ 3,,,,.j, 2
00
/ I ,,,,,,,EDS ABORT LIMIT
I i_GINE ,NO.5
T''.4-.=....._G,NE,No.1
ENGINE NO. 2
,IL10 20 30 40 50 60 70
TIME OF ENGINE FAILURE (SEC)
Figure 2-45
CH.2011_
S-IVBPEAKPITCHATTITUDEERRORFORS-IB ENGINE1, 2, & 5 FAILURES
14
0== 12
00 10
ENGINE NO, 5
ENGINE NO. 1
ENGINE NO. 2
\ENGINE
20 30 40 50
TIME OF ENGINE FAILURE (SEC)
60 70
Figure 2-44
2.41
2.0
G_'_ 1.6
_ ,._
a ZO
.4
0
8
_u
O
f
i'
IENGINE NO. 5
I II I
ENGINE NO. 1
[I_,_,_,_ _ = J-2 GIMBAL LIMIT
ENGINE NO. I _
20 30 40 50 60
TIME OF ENGINE FAILURE (SEC)
Figure 2-46
2-15
SectionIIPerformance
MASS CHARACTERISTICS.
The following mass characteristics data were extracted from the
latest final predicted mass characteristics studies, which are based
on measured weights of the dry stages and hardware and propulsion
systems performance predictions. Figure 2-47 lists the vehicle weight
breakdown at liftoff. Figures 2-48 and 2-49 contain a weight break-
down at major events during S-IB and S-IVB powered flight,
respectively. Figure 2-50 shows the total residuals at J-2 enginecutoff command.
St-2 VEHICLE WEIGHT BREAKDOWN (LBM)
CM 13,500
SM 17,500
SLA PANELS 2,800
SLA (FIXED) 1,500
INSTRUMENT UNIT 4,298
S-IVB STAGE INERT 24,465
USEABLE S-IVB PROPELLANT (INCLUDES FPR) 3,727ORBIT INSERTION WEIGHT
J-2 THRUST DECAY PROPELLANT 121S-IVB CUTOFF WEIGHT
S-IV8 PROPELLANT CONSUMED 227,637
S-IVB APS PROPELLANT CONSUMED 6
LES 9,350ULLAGE CASES 214
S-IVB "90% THRUST" WEIGHT
S-IVB GH2 START TANK 4
S-IVB BUILDUP PROPELLANT CONSUMED 565
ULLAGE PROPELLANT CONSUMED 176
S-IVB WEIGHT AT SEPARATION
S-IVS AFT FRAME HARDWARE 31
S-IB,/S-iVB INTE RSTAGE 6,800
S-IB DRY WEIGHT 84,600
S-IB RESIDUALS AND RESERVES 10,718
S-IVB DETONATION PACKAGE 5
S-IVB FROST CONSUMED 100
S-IS FROST CONSUMED 1,000
S-IB SEAL PURGE CONSUMED 6
S-IB FUEL ADDITIVE CONSUMED 27
S-IB GEARBOX CONSUMPTION (RP-1) 706
INBOARD ENGINE THRUST DECAY FRPT CONSUMED 2,162OUTBOARD ENGINE THRUST DECAY PRPT CONSUMED
TO SEPARATION 1,672
S-IS MAINSTAGE PROPELLANT CONSUMED 882,092VEHICLE LIFTOFF WEIGHT
Figure 2-47
67,790
67,911
305,118
305,863
1,295,782
CH.20114
DESCRIPTION GROUND ING FIRST MOTION IECO SIGNAL OECO SIGNAL SEP SIGNAL SEP COMP
S-IB DRY
LOX IN TANKS
LOX BELOW TANKS
LOX ULLAGE (GOX)
FUEL IN TANKS
FUEL BELOW TANKS
FUEL ULLAGE (He)
FUEL PRESS. HELIUM SUPPLY
GN2
ORONITE
HYDRAULIC OIL
ICE
S-IVB AFT FRAME
TOTAL S-IB
S-IB/S-IVB INTERSTAGERETRO-PROPELLANT
TOTAL S-I8/S-IVB INTERSTAGE
FIRST VEHICLE STAGE
S-IVB LOADED
VIU
LES
CM
SERVICE MODULE
SLA WITH RING
TOTAL VEHICLE
2-16
84,600
624,033
7,760
32
274,833
4,826
5
78
15
33
28
I0_0
0
84,600612,549
8,201
80
270,637
5,7588
75
15
33
28
1000
0
84,600
2,345
8,094
2,619
4,794
5,758
58
25
9
6
28
0
0
84,600
03,317
2,651
1,004
5,292
59
24
9
628
0
0
84,6000
2,767
2,653
122
5,08059
24
9
6
28
0
0
997,243
5,738
1,062
6,800
1,004,043
257,051
4,298
9,350
13,500
17,500
4,300
1,310,042
982,984
S,738
1,062
6,800
108,336
5,738
1,062
6,800
96, 990
5,738
1,062
6,800
95,348
5,738
1,062
6,800989,784
257,051
4,298
9,350
13,500
17,500
4,300
1,295,783
115,136
256,951
4,298
9,350
13,500
17,500
4,300
421,035
103,790
256,951
4,298
9,350
13,500
17,500
4,300
409,689
102,148
256,951
4,298
9,350
13,500
17,500
4,300
408,047
84,6000
2,761
2,653115
5,064
59
24
9
6
28
031
95,350
5,738
1,062
6,800
102,150
102,150
Figure 2-48
ETD
84,600
0
2,683
2,652
111
4,827
59
24
9
6
28
0
31
95,030
5,738
1,062
6,800
101,830
101,830
Section II Performance
TRACKING COVERAGE.
Tracking, command, and communications systems, and telemetry
coverage data were extracted from the latest launch vehicle opera-tional flight trajectory tracking analysis, for the SA-207 vehicle.
During the launch phase, overlapping coverage is provided by sites
at MILA, Bermuda and Newfoundland. Figure 2-51 shows a map
trace of this launch phase coverage. Figure 2-52 contains a summaryof orbital tracking coverage of the S-IVB/IU with orbital traces
on a world map. The tracking and communications network shown
on this figure is currently planned to be available for Saturn IB
flights in the Skylab program.
POWERED FLIGHT TRACKING COVERAGE (TYPICAL
/_ _iiii!ii!:_:<'_ii::i::i_:_.................iii_i::::iii!::i::iiii:_._:,_<_,.......
_ ..<..< ""b_ ::::::"" ,.
,.:_i::i:iiii,_,....... J
_- &=TINSE=,ON
30 o
-90 ° -75 ° -60 °
LONGITUDE
Figure 2-51
-45 °
CH-20118
ii
-100 ° .80 °
(2-,_,g
.100 °
-60 °
.80 °
LONGITUDE
-40 ° -20 o 0 o 20° 40 ° 60 ° 80° 100 ° 120 ° 140° 160 ° *-180 ° .160 o -140 ° .120 ° -I00O -_0o
.60 o .40 ° .20 ° 0 o 20° 400 60° 80° I00 ° 120 ° 140 ° 160° _+180° .160 ° .140 ° -120 ° -I00 o .80 °
STATIONS:
I. M_LA 4. CANARY ISLAND 7. CARNARVON I0, HAWAII
2. BERMUDA 5. ASCENSION 8. HONEYSUCKLE 11. GOLDSTONE
3. NEWFOUNDLAND 6. MADRID 9. GUAM 12. CORPUS CHRISTI
NOTE: GROUND STATION VISIBILITY CONTOURS REFLECT KEYHOLE AND TERRAIN LIMITATIONS
80 °
60o
400
20°
0o
20°
40o
_0o
CH-20119
2-18
Figure 2-52
TABLEOFCONTENTS
LaunchVehicleMonitoringAndControl...................... 3-1
Launch Vehicle Monitoring Displays ............................ 3-1Launch Vehicle Normal Controls....; ............................. 3-10
Launch Vehicle EDS Controls ...................................... 3-11
Abort Controls ................................................................ 3-12
Abort Modes And Limits ................................................ 3-15
LAUNCH VEHICLE MONITORINGAND CONTROL.
The spacecraft is equipped with a number of displays and controls
which permit monitoring the launch vehicle conditions and control
ling the launch vehicle under normal and emergency conditions.
Many of these displays and controls are related to the Emergency
Detection System (EDS). The displays implemented for EDS moni-
toring were selected to present as near as possible those parameters
which represent the failures leading to vehicle abort. Whenever
possible, the parameter was selected so that it would display totalsubsystem operation. Manual abort parameters have been imple-
mented with redundant sensing and display to provide highly
reliable indications to the crewmen. Automatic abort parameters
have been implemented triple redundant, voted two-out-of-three,
to preclude single point hardware or sensing failures causing an
inadvertent abort. The types of displays have been designed to
provide onboard detection capability for rapid rate malfunctions
which may require abort. Pilot abort action must, in all cases, bebased on two separate but related abort cues. These cues may be
derived from the EDS displays, ground information, physiological
cues, or any combination of two valid cues. In the event of a
discrepancy between onboard and ground based instrumentation,onboard data will be used. The EDS displays and controls are
shown in figure 3-1. As each is discussed it is identified by use
of the grid designators listed on the border of the figure.
LAUNCH VEHICLEMONITORING DISPLAYS.
FLIGHT DIRECTOR ATTITUDE INDICATOR.
There are two Flight Director Attitude Indicators (FDAI's), each
of which provides a display of Euler attitude, attitude errors and
angular rates. Refer to figure 3-1, Q-45 and J-60, for locationsof the FDAI's on the MDC and to figure 3-2 for details of the
FDAI's. These displays are active at liftoff and remain active
throughout the mission, except that attitude errors are not displayed
during S-I VB flight. The FDAI's are used to monitor normal launch
vehicle guidance and control events. The roll and pitch programsare initiated simultaneously at + I0 seconds. The roll program is
terminated when flight azimuth is reached, and the pitch programcontinues to tilt-arrest. IGM initiate will occur approximately one
second after LET jettison during the S-IVB stage flight. The FDA!
ball displays Euler attitude, while needle type pointers across the
face of the ball indicate attitude errors, and triangular pointers
around the periphery of the ball display angular rates. Attitude
errors and angular rate displays are, clockwise from the top, roll,
pitch and yaw, respectively. Signal inputs to the FDAI's are switchselectable and can come from a number of different sources in
the spacecraft. This flexibility and redundancy provides the required
attitude and error backup display capability. Excessive pitch, roll,
or yaw indications provide a single cue that an abort is required
Additional abort cues will be provided by the FDAI combiningrates, error, or total attitude. Second cues will also be provided
by the LV RATE light (R-50), LV GUID light (R-52), physiologicalsensations and MCC ground reports.
LV ENGINES LIGHTS.
The eight LV ENGINES lights on (S-5I, figure 3-1) indicate
that each corresponding S-IB stage engine is below 90 percentnominal thrust. The engine light cluster also provides indications
of the launch vehicle staging sequence. Physical separation
of stages is indicated with all engine lights off (after normal S-IB
cutoff). The no. 1 light will come on again 2.4 sec after OECO
and go off when the S-IVB engine exceeds 65 percent nominalthrust. For abort decisions, the on indication is considered zero
thrust for the corresponding engine and off is 100 percent thrust.
Each S-IB engine light and sensing circuit is redundant and consti-
tutes a single "warning" cue of a possible abort situation. However,
it is not an abort cue in itself. Upon ground verification of a single
engine failure, the ABORT SYSTEM 2 ENGINE OUT switchshould be moved to OFF (if T > + 15 sec). A second engine
light is not necessarily a second cue for immediate abort. For thisfailure case (two non-adjacent engine lights on), secondary abort
indications will be provided by the FDAI, LV RATE light, and
ground information. The Saturn IB vehicle has limited capabilityto continue the mission with the loss of one or two engines depend-
ing upon the time of the second failure. Mission rules will prescribe
the timelines after which flight will be continued with one or two
engines out. Simultaneous illumination of two or more engine lights
at any time during S-1B flight is sut_cient cue for immediate abort
except at normal IECO or OECO. During S-IVB burn the no.
1 engine light on is a single abort cue.
Two engine out automatic aborts are active until manually deac-tivated by the crew. Deactivation times will be prescribed by mission
rules. Two engine automatic abort disabling by the flight crew
is backed up by the launch vehicle sequencer prior to IECO and
is not reactivated.
LV RATE LIGHT.
The LV RATE light on (R-50, figure 3-1) is the primary cue from
the launch vehicle that preset overrate settings have been exceeded.
It is a single cue for abort, while secondary cues will be provided
by FDAI indications, physiological cues, or ground information.Automatic LV rate aborts are enabled automatically at liftoff (with
EDS and LV RATES AUTO switches enabled in SC) and are
active until deactivated by the crew. EDS auto abort deactivation
times will be governed by mission rules. The automatic LV rate
abort capability is also deactivated by the launch vehicle sequencer
3-1
[
Section III Emergency Detection and Procedures
MAIN DISPLAY CONSOLE
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Section III Emergency Detection and Procedures
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Section III Emergency Detection and Procedures
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125 126 127 128 129 130 131 132 133 134 135 136 137 138 139 140 141 142 143 144 145 146 147
Figure 3-1 (Sheet 6 of 7)
148 149 150 151
CH.20137
3-7
Section III Emergency Detection and Procedures
ABORT LIGHT N-51 iMU CAGE SWITCH AE-36
ABORT SYSTEM SWITCHES AA-59 LES MOTOR RRE PUSHBUTTON X-49
ACCELEROMETER O-40 LI ETO FF/NO AUTO ABORT LI GHTS W-50
ALTIMATER C-51 LIMIT CYCLE SWITCH T-37
APEX COVER JETT PUSHBUTTON W-52 LOGIC POWER SWITCH AB-13
ATTITUDE SET CONTROL PANEL AD-32 LV ENGINE LIGHTS S-51
ATT DEADBAND SWITCH T-38 LV GUID LIGHT R-52
ATT SET SWITCH R-40 LV RATE LIGHT R-50
ATVC GAI N SWITCH AC-46 LV RATES SWITCH AA-59
AUTO RCS SELECT SWITCHES 0-14 LV/SPS IND SWITCHES AE-40
BMAG MODE SWITCHES Z-32 LV TANK PRESS GAUGES X-46
BMAG PWR SWITC_IES AA-9 LV_/SPS Pc INDICATOR W-42
CANARD DEPLOY PUSHBUTTON Y-49 MAIN DEPLOY PUSHBUTTON Y-52
CMC ATT SWITCH 0-37 MAIN RELEASE SWITCH AG-56
CMC MODE SWITCH W-39 MANUAL ATTITUDE SWITCHES T-34
CM PRPLNT SWITCHES AC-51 MASTER ALARM LIGHT L-40,0-110
CM RCS LOGIC SWITCH AC-50 PRPLNT DUMP SWITCH AA-56
CM RCS He DUMP PUSHBLITTON Z-52 RATE SWITCH T-39
CM RCS PRESS SWITCH 0-68 REACTION CONTROL SYSTEM C,/B U-8
CM/SM SEP SWITCHES Y-61 ROLL STABIUTY INDICATOR L-45
CORRIDOR INDICATORS K-4S, L-45 ROT-CONTR PWR SWITCHES W-33
CSM/LV SEP PUSHBUTTON Zo49 SC CONT SWITCH W-38
DIRECT 0 2 SWITCH AD-10 SCS ELECTRONICS POWER SWITCH AA-10
DIRECT ULLAGE PUSHBUTTON Z-38 SCS TVC SERVO POWER SWITCHES AA-15
DOCK RING SEP SWITCHES Y-59 SCS TVC SWITCHES AC-39
DROGUE DEPLOY PUSHBUTTON X-52 SERVICE PROPULSION SYSTEM C/B W-10
DSKY PANEL R-60 SIG COND/DRIVER BIAS POWER SWITCHES AC-12
EDS BATTERY C/B X-5 SPS THRUST LIGHT K-49
EDS POWER SWITCH Z-17 SPS THRUST SWITCH Z-36
EDS AUTO SWITCH Y-S7 SPS GIMBAL THUMBWHEEL CONTROLS AA-46
ELS/CM-SM SEP C/B X-9 SPS GIMBAL MOTORS SWITCHES AC-42
ELS SWITCHES AC-48 STABILIZATION CONTL SYS C/B P-S,R-10, S-9
ENLS FUNCTION SWITCH G-45 THRESHOLD INDICATOR (.05 G LIGHT) K-48
EMS MODE SWITCH 1-44 THRUST ON PUSHBUTTON AA-38
ENTRY SWITCHES AE-38 TRANS CONTR SWITCH T-40
EVENT TIMER INDICATOR P°51 TVC GIMBAL DRIVE SWITCHES AE-42
EVENT TIN_R SWITCHES AEoSO TWR JETT SWITCHES AA-61
FDAI J-60,Q-45 UP TLM SWITCH 0-66
FDAI CONTROL SWITCHES R-37 2 ENG OUT SWITCH AA-58
FDAI/GPI POWER SWITCH AA-12 L_V THRUST SWITCHES Z-41
GDC ALIGN PUSHBUTTON AC-36 _V/EMS SET SWITCH K°52
GUIDANCE SWITCH AC-56 /W/RANGE COUNTER L-48
G-V PLOTTER H-49
NOTE:
THE PANEL INDICATORS AND SWITCHES ASSOCIATED WITH THE EDS AND LAUNCH VEHICLE CONTROL ARE LISTED
FOR EASE IN LOCATION. REFER TO THE TEXT FOR DISCUSSION OF SWITCH AND INDICATOR FUNCTIONS.
CH-20131
3-8
Figure 3-1 (Sheet 7 of 7)
prior to IECO and is not active during S-IVB flight. After the
automatic abort system is deactivated the LV RATE light is used
only to indicate that preset LV overrate settings have been exceeded.
The light is a single cue for abort. Secondary cues will be provided
by the FDAI, physiological cues, or ground communications, The
LV preset overrate settings are:
Pitch 5 ___ 0.5 deg/sec Liftoff to J-2
and Yaw ignition
l0 -+- 0.5 deg/sec J-2 ignition toJ-2 cutoff
Roll 20 -----0.5 deg/sec Liftoff to J-2cutoff
The LV RATE light will come on at any time during first or second
stage flight if the LV rates exceed these values.
Note
The LV RATE light may blink on and offduring normal
staging.
LV GUID LIGHT.
The LV platform (ST-124M-3) is interrogated every 40 ms for the
correct attitude. Three or more excessive attitude discrepancy read-
ings per second, in any one axis, will cause the system to switch
to the coarse resolvers. Fifteen or more excessive attitude discrep-
ancy readings per second on the coarse resolvers, in any one axis,
will inhibit the LV attitude change commands being sent to the
flight control system. The flight control system will then hold the
last acceptable command.
A signal is sent from the LVDA to activate the LV GUID light
(R-52, figure 3-1) at the same time the flight control commands
are inhibited. It is a single cue for abort. Second cues will be
Section III Emergency Detection and Procedures
provided by the LV RATE light (only when the automatic abort
system is on) and by the FDAI, ground information, or both.
The LV GUID light ON is a prerequisite to spacecraft takeover
of the Saturn during launch vehicle burn phases. See subsequent
paragraph on GUIDANCE SWITCH for a further discussion of
spacecraft takeover.
LIFTOFFINO AUTO ABORT LIGHTS.
The LIFTOFF and NO AUTO ABORT lights (W-50, figure 3-1)
are independent indications contained in one switch/light assembly.
The LIFTOFF light ON indicates that vehicle release has been
commanded and that the IU umbilical has ejected. The spacecraft
digital event timer is started by the same function. The LIFTOFF
light is turned OFF at S-IB IECO.
The NO AUTO ABORT light ON indicates that one or both of
the spacecraft sequencers did not enable automatic abort capability
at liftoff. Automatic abort capability can be enabled by pressing
the switch/light pushbutton. If the light remains ON one or both
of the automatic abort circuits failed to energize. The crew must
then be prepared to back up the automatic abort manually. The
NO AUTO ABORT light is also turned OFF at S-IB IECO.
WARNING
If the NO AUTO ABORT pushbutton is depressed at
T-0 and a pad shutdown occurs, a pad abort will result.
ABORT LIGHT.
The ABORT light (N-51, figure 3-1) can be illuminated by ground
command from the Flight Director, the Mission Control Center
(MCC) Booster Systems Engineer, the Flight Dynamics Officer,
the Complex 39 Launch Operations Manager (until tower clearance
FLIGHTDIRECTORATTITUDEINDICATORROLL
+ANGULAR VELOCITY-
+ATTITUDE ERROR-
NOTE:
ALL POLARITIES INDICATE
VEHICLE DYNAMICS
PITCH & YAW
EULER ATTITUDE ON BALL
YAW - _1 = 034 °
ROLL - _ = 330 °
ATTrTUDE
ERROR
ANGULAR
T VELOCITYC
+
SCALE SELECT SOURCE
FDAI RATE & DISPLAY SELECT SWITCHES
Figure 3-2
YAW
+ATTITUDE ERROR-
+ANGULAR VELOCITY-
OTAL
ATTITUDE SCALE
3-9
Section III Emergency Detection and Procedures
at + 10 see), or in conjunction with range safety booster engine
cutoff. The ABORT light ON constitutes one abort cue. An RF
voice abort request constitutes one abort cue.
Note
Pilot abort action is required prior to receipt of an
ABORT light or a voice command for a large percentage
of the time critical launch vehicle malfunctions, particu-
larly at liftoff and staging.
ANGLE OF ATTACK METER.
The ar/gle of attack (Qa) meter (W-42, figure 3-1) is time shared
with service propulsion system (SPS) chamber pressure. The Qa
display is a pitch and yaw vector summed angle-of-attack/dynamic
pressure product (Qa). It is expressed in percentage of total pressure
for predicted launch vehicle breakup (breakup limit equals 100%).It is effective as an information parameter only during the high.
Q flight region from + 50 see to + 1 min 40 sec. Except as statedabove, during ascent, the Qa meter provides trend information
on launch vehicle flight performance and provides a secondary
cue for slow-rate guidance and control malfunctions. Primary cues
for guidance and control malfunctions will be provided by the
FDAI, physiological cues, and/or MCC callout.
Nominal angle of attack meter indications should not exceed 25%.
Expected values based on actual winds aloft will be provided by
MCC prior to launch.
ACCELEROMETER.
The accelerometer (0-40, figure 3-1) indicates longitudinal acceler-
ation/deceleration. It provides a secondary cue for certain engine
failures and is a gross indication of launch vehicle performance.
The accelerometer also provides a readout of G-forces during
reentry.
ALTIMETER.
Due to dynamic pressure, static source location, and instrument
error the altimeter (C-51, figure 3-1) is not considered to be an
accurate instrument during the launch phase. The primary function
of the altimeter is to provide an adjustable reference (set for baro-
metric pressure on launch date) for parachute deployment for
pad/near pad LES aborts. However, the aerodynamic shape of
the CM coupled with the static source location produces errors
up to 1800 feet. Therefore, the main parachutes must be deployed
at an indicated 3800 feet (depends on launch day setting) to ensure
deployment above 2000 feet true altitude.
EVENT TIMER.
The event timer (P-51, figure 3-1) is a critical display because it
is the primary cue for the transition of abort modes, manual se-
quenced events, monitoring roll and pitch program, staging, andS-IVB insertion cutoff. The event timer is started by the liftoff
command which enables automatic aborts. The command pilot
should be prepared to manually back up its start to ensure timer
operation. The event timer is reset to zero automatically with abortinitiation.
MASTER ALARM LIGHT.
There are three MASTER ALARM lights, one on main display
panel 1 (L-40, figure 3-1), one on main display panel 3 (O-110)
and one in the lower equipment bay. The three MASTER ALARM
lights ON alert the flight crew to critical spacecraft failures orout-of-tolerance conditions identified in the caution and warning
light array. After extinguishing the alarm lights, action should be
3-10
initiated to correct the failed or out-of-tolerance subsystem. If crew
remedial action does not correct the affected subsystem, then an
abort decision must be made based on other contingencies. Secon-
dary abort cues will come from subsystem displays, ground verifi-
cation, and physiological indications.
Note
The Commander's MASTER ALARM light (L-40) will
not illuminate during the launch phase, but the other
two MASTER ALARM lights can illuminate and thealarm tone will sound.
LV TANK PRESS GAUGES.
The LV TANK PRESS Gauges (X-46, figure 3-1) indicate the S-IVB
tank pressures. The two left-hand pointers indicate S-IVB oxidizer
pressure. The two right-hand pointers indicate S-IVB fuel tank
pressure until LV/spacecraft separation.
LAUNCH VEHICLENORMAL CONTROLS.
GUIDANCE SWITCH.
The GUIDANCE switch is a two position guarded toggle switch
with the two positions being IU and CMC (AC-56, figure 3-1).
The switch controls a relay in the IU which selects either the IU
or the CMC in the spacecraft as the source of flight control attitude
error signals for the LV. The normal position of the GUIDANCE
switch is IU. Placing the switch in the CMC position permits
spacecraft control of the LV under certain conditions.
Guidance Reference Failure Condition.
During the LV burn modes, when the LVDC recognizes a guidancereference failure and turns on the LV GUID light, the GUIDANCE
switch can be placed in the CMC position. (The switch function
is interlocked in the LVDC such that a guidance reference failure
must be recognized before the CMC switch position will be hon-
ored.) With the switch in the CMC position and LVDC recognition
of a guidance reference failure, the LV will receive attitude error
signals from the CMC via the LVDC. During the S-IB burn phase
the CMC provides attitude error signals based on preprogrammed
polynomial data. During S-IVB burn phases the rotational hand
control (RHC) is used to generate the attitude error signals. Theamount of attitude error transmitted to the LVDC is a function
of how long the RHC is out of detent. The RHC in this mode
is not a proportional control. S-IVB engine cutoff must be executedwith the RHC ccw (_ 3 see) based on observation of the computed
velocity display.
No Guidance Reference Failure Condition.
During the coast mode, T4, the guidance reference failure is notinterlocked with the GUIDANCE switch and the spacecraft can
assume control of the LV any time the switch is placed in the
CMC position. With the GUIDANCE switch in the CMC position
and no guidance reference failure, the LVDC will function in a
follow-up mode. When LV control is returned to the LVDC during
T 4 attitude orientation will be maintained with reference to localhorizontal.
EDS POWER SWITCH.
The EDS POWER switch (Z-17, figure 3-1) should be in the EDs
power position during prelaunch and launch operations. The switch,
ifplacedintheOFFpositionresultsinan"EDSUnsafe"function.Thisfunctionhasthefollowingeffectonthelaunchcountdown.1If theswitchisturnedOFFbeforeautomaticsequence,theautomaticsequencewillnotbeentered.2.If theswitchisturnedOFFduringtheautomaticsequence(startingatT-187see),butpriortoT-30.0sec,thecountdownisstoppedatT-30.0secandthecountdownisrecycledtoT-24min.3.If theswitchisturnedOFFafterT-30sec,butpriortoT-16.2sec,thecountdownisstoppedatT-16.2secandthecountdownisrecycledtoT-24min.4.If theswitchisturnedOFFafterT-16.2sec,butpriortoT-3.1sec(ignitioncommand),thecountdownisstoppedatT-3.1secandthelaunchisrecycledtoT-24min(orisscrubbedforthatday).5.If theswitchisturnedOFFafterT-3.1sec,butpriortoT-50ms(launchcommit),thecountdownisstoppedimmediatelyandthelaunchisscrubbedforthatday.6.AfterT-50ms,theswitchwillnotstoplaunch;however,theSCEDSpoweroffwillhavethefollowingeffectonthemission.a.TheEDSdisplayswillnotbeoperative.b.TheAuto-Abortcapabilitywillnotbeenabled;however,manualabortcanbeinitiated.
WARNING
If the SC EDS POWER switch is returned to the ON
position after liftoff, an immediate abort may result,
depending upon which relays in the EDS circuit activatefirst.
LAUNCH VEHICLE EDS CONTROLS.
EDS SWITCH.
The EDS switch is a two position toggle switch with the two posi-
tions being AUTO and OFF (Y-57, figure 3-1). Prior to liftoff
the EDS switch is placed in the AUTO position so that an automatic
abort will be initiated if:
1. A LV structural failure occurs between the IU and the CSM.
2. Two or more S-IB engines drop below 90% of rated thrust.
3. LV rates exceed 5 degrees per second in pitch or yaw or 20
degrees per second in roll.
The two engine out and LV rate portions of the auto abort system
can be manually disabled, individually, by the crew (Normally atT+ 1 min 40 see). The LV RATES and the 2 engine out is automat-
ically disabled just prior to IECO.
DOCK RING SEP SWITCHES.
The DOCK RING SEP switches are a pair of two position guarded
toggle switches (Y-59, figure 3-1). Their purpose is to provide ameans of manually initiating final separation of the LM docking
ring. During a normal entry or an SPS abort, the docking ring
must be jettisoned by actuation of the DOCK RING SEP switches.
Failure to jettison the ring could possibly hamper normal earth
landing system (ELS) functions.
CM/SM SEP SWITCHES.
The two CM/SM SEP switches (Y-6I, figure 3-1) are redundant,
Section III Emergency Detection and Procedures
momentary ON, guarded switches, spring loaded to the OFF posi-
tion. They are normally used by the command pilot to accomplish
CM/SM separation prior to the reentry phase. These switches canalso be used to initiate an LES abort in case of a failure in either
the EDS or the translational controller. All normal post-abort events
will then proceed automatically. However, the CANARD DEPLOY
pushbutton (Y-49, figure 3-1) should be depressed 11 seconds after
abort initiation, because canard deployment will not occur if thefailure was in the EDS instead of the translational controller.
PRPLNT SWITCH.
The PRPLNT switch is a two position toggle switch with the two
positions being DUMP AUTO and RCS CMD (AA-56, figure 3-1).
The switch is normally in the DUMP AUTO position prior to liftoff
in order to automatically dump the CM reaction control system
(RCS) propellants, and fire the pitch control (PC) motor if an abortis initiated during the first 61 seconds of the mission. The propellant
dump and PC motor are inhibited by the SC sequencer at 61 sec.The switch in the RCS CMD position will inhibit propellant dump
and PC motor firing at any time.
ABORT SYSTEM--2 ENG OUT SWITCH.
The 2 ENG OUT switch is a two position toggle switch, the two
positions being AUTO and OFF (AA-58, figure 3-1). The purposeof this switch is to enable or disable EDS automatic abort capability
for a two engine out condition. Normal position of the switch isAUTO, which enables the EDS automatic abort capability. With
the switch in OFF the EDS automatic abort capability is disabled.
ABORT SYSTEM--LV RATES SWITCH.
The LV RATES switch is a two position toggle switch, the two
positions being AUTO and OFF (AA-59, figure 3-1). The purposeof this switch is to enable or disable EDS automatic abort capability
for excessive LV rates. Normal position of the switch is AUTO,
which enables the EDS automatic abort capability for excessive
LV rates. Placing the switch in OFF disables the capability. The
capability is disabled automatically just prior to IECO.
ABORT SYSTEM--TWR JETT SWITCHES.
There are two redundant TWR JETT guarded toggle switches
(AA-61, figure 3-1). When these switches are placed in AUTO,
explosive bolts and the tower jettison motor are fired to jettison
the LET. Appropriate relays are also de-energized so that if anabort is commanded, the SPS abort sequence and not the LES
sequence will occur.
MAIN RELEASE SWITCH.
The MAIN RELEASE switch (AG-56, figure 3-1) is a toggle switch
guarded to the down position. It is moved to the up position to
manually release the main chutes after the command module has
landed. No automatic backup is provided. This switch is armed
by the ELS LOGIC switch ON and the 10K barometric switches
closed (below 10,000 feet altitude).
Note
The ELS AUTO switch must be in the AUTO position
to allow the 14-see timer to expire before the MAIN
CHUTE RELEASE switch will operate.
ELS SWITCHES.
There are two ELS two position toggle switches (AC-48, figure
3-1). The left hand switch is guarded to the OFF position and
should only be placed in the LOG IC position during normal reentry
3-1 1
Section III Emergency Detection and Procedures
or following an SPS abort, and then only below 45,000 feet altitude.
If the ELS LOGIC and AUTO switches are activated at any time
below 24,000 feet (pressure altitude), the landing sequence will
commence, i.e., LES and apex cover jettison and drogue deploy-ment. If activated below 10,000 feet altitude, the main chutes will
also deploy. ELS LOGIC is automatically enabled following any
manual or auto EDS initiated LES abort. It should be manuallybacked up if time permits.
WARNING
Do not use ELS LOGIC and ELS AUTO switches
during normal launch. Activation of ELS LOGIC and
ELS AUTO switches below 40,500 feet during ascent
will initiate the landing sequence causing LES and apex
cover jettison and deployment of drogue chutes.
The right hand switch is not guarded and has positions of AUTO
and MAN. Its normal position is MAN until AUTO is required
at + 14 sec on mode IA and IB aborts or at 30,000 ft on high
altitude aborts or entry to enable the automatic sequencing of the
ELS during a CM descent period. If the switch is placed in MAN
it will inhibit all automatic sequencing of the ELS.
CM RCS PRESS SWITCH.
The CM RCS PRESS switch is a two position guarded toggle switch(0-68, figure 3-1). Any time the CM is to be separated from the
SM, the CM RCS must be pressurized. The normal sequence of
events for an abort or normal CM/SM SEP is to automatically
deadface the umbilicals, pressurize the CM RCS, and then separate
the CM/SM. However, if the automatic pressurization fails, theCM RCS can be pressurized by the use of the CM RCS PRESSswitch.
ABORT CONTROLS.
TRANSLATIONAL CONTROLLER.
The TRANSLATIONAL CONTROLLER, which is mounted on
the left arm of the commanders couch, can be used to accomplish
several functions. A manual LES abort sequence is initiated by
rotating the T-handle fully ccw. This sends redundant engine cutoff
commands to the LV (engine cutoff from the SC is inhibited during
the first 30 sec of flight), initiates CM/SM separation, fires the
LES motors, resets the SC sequencer and initiates the post abortsequence. For a manually initiated SPS abort, the ccw rotation
of the T-handle commands LV engine cutoff, resets the SC se-
quencer and initiates the CSM/LV separation sequence.
Note
Returning the T-handle to neutral before the 3 sec
expires results only in an engines cutoff signal ratherthan a full abort sequence.
CW rotation of the T-handle transfers control of the SC from the
CMC to the SCS. The T-handle can also provide translation control
of the CSM along one or more axes. The T-handle is mounted
approximately parallel to the SC axis; therefore, T-handle move-
ment will cause corresponding SC translation. Translation in the
+ X axis can also be accomplished by use of the direct ullage
pushbutton; however, rate damping is not available when usingthis method.
SEQUENCER EVENTS--MANUAL PUSHBU'rTONS.
These are a group of covered pushbutton switches (X-51, figure
3-12
3-1) which provide a means of manual backup for abort and normal
reentry events which are otherwise sequenced automatically.
LES MOTOR FIRE Switch.
The LES MOTOR FIRE switch is used to fire the launch escape
motor for an LES abort if the motor does not fire automatically.
It is also a backup switch to fire the LET jettison motor in the
event the TWR JETT switches fail to ignite the motor.
CANARD DEPLOY Switch.
The CANARD DEPLOY switch is used to deploy the canard in
the event it does not deploy automatically during an abort.
CSM/LV SEP Switch.
The CSM/LV SEP switch is used as the primary means of initiatingCSM/LV separation after the ascent phase of the mission. When
the switch is pressed it initiates ordnance devices which explosivelysever the SLA, circumferentially around the forward end, and
longitudinally, into four panels. The four panels are then rotated
away from the LV by ordnance thrusters. Upon reaching an angle
of 45 degrees spring thrusters jettison the panels away from theSC. The same ordnance train separates the CSM/LV umbilical.
The CSM/LV SEP switch is also used as a backup to initiateseparation of the SLA when an SPS abort cannot be initiated from
the TRANSLATIONAL CONTROLLER. The +X translation
would have to be manually initiated under these curcumstances.
APEX COVER JETT Switch.
The APEX COVER JETT switch is used to jettison the APEX
COVER in the event it fails to jettison automatically during anabort or a normal reentry.
DROGUE DEPLOY Switch.
The DROGUE DEPLOY switch is used to deploy the drogue
parachutes in the event they fail to deploy automatically 2 secafter the 24,000-foot barometric pressure switches close.
MAIN DEPLOY Switch.
The MAIN DEPLOY switch is used to deploy main parachutes
in the event they fail to deploy automatically when the 10,000-footbarometric pressure switches close. This switch can also be used
to manually deploy the main parachutes during mode IA aborts.
CM RCS He DUMP Switch.
The CM RCS He DUMP switch is used to initiate depletion of
the CM He supply if depletion does not occur normally as an
automatic function during abort.
SERVICE PROPULSION SYSTEM (SPS) CONTROL.
The SPS provides primary thrust for major velocity changes sub-
sequent to SC/LV separation and prior to CM/SM separation.The SPS is also used to accomplish mode III and IV aborts.
SPS Engine Start.
SPS engine ignition can be commanded under control of the CMC,
the SCS or manually. For all modes of operation, the AV THRUST
switch A (Z-40), or _V THRUST switch B (Z-42), or both, must
be in the NORMAL position. (If double-bank operation is desired.A V THRUST switch B is moved to NORMAL 5 seconds or more
following SPS ignition.) Ullage is normally provided by the THC
(+ X translation) and backup is by DIRECT ULLAGE pushbutton(Z-38, figure 3-1). The DIRECT ULLAGE pushbutton is a momen-
tary switch and must be held depressed until the ullage maneuver
iscomplete.Itdoesnotprovideratedamping.TheSPSTHRUSTlight(K-49)willilluminatewhentheengineisfiring.IntheCMCmodeTHRUSTONiscommandedasaresultofinternalcomputa-tions.PrerequisitesareSCCONTswitch(W-38)inCMCpositionandTHCinneutral(exceptfor+ X translation). In the SCS mode,
the SC CONT switch must be in the SCS position, or the THC
rotated cw. SPS ignition is commanded by pressing the THRUST
ON pushbutton (AA-38). Prerequisites are +X translation (fromTHC or DIRECT ULLAGE pushbutton) and AV/RANGE
counter _> 0 (L-48). Once SPS ignition has occurred in this mode,+X translation and THRUST ON commands can be removed
and ignition is maintained until a THRUST OFF command is
generated. In the manual mode the SPS THRUST switch (Z-36)
is placed in DIRECT ON. Ignition is maintained until a THRUST
OFF command is generated.
WARNING
The SPS THRUST switch is a single-point failure with
the AV THRUST switches in the NORMAL position.
SPS Engine Shutdown.
In the CMC mode, normal engine shutdown is commanded by
the CMC as a result of internal computations. In the SPS mode,
engine shutdown can be commanded by the EMS ±V counter
running down to 0 or by placing the AV THRUST switches (both)
to OFF. In the manual mode, shutdown is commanded by placingthe AV THRUST switches (both) to OFF.
Thrust Vector Control.
Four gimbal motors control the SPS engine position in the pitch
and yaw planes; two motors for each plane. These motors are
activated by the SPS GIMBAL MOTORS switches (AC-42, figure
3-1).
Note
The motors should be activated one at a time due to
high current drain during the start process.
Control signals to the gimbal motors can come from the CMC,SCS or the RHC. Gimbal trim thumbwheels (AA-46) can also be
used to position the gimbals in the SCS ,SV mode. The TVCGIMBAL DRIVE switches (AE-42) are three position toggle
switches. Their purpose is to select the source and routing of TVC
signals. The switches are normally in the AUTO position.
STABILITY CONTROL SYSTEM (SCS).
The SCS is a backup system to the primary guidance navigation
and control system (PGNCS). It has the capability of controlling
rotation, translation, SPS thrust vector and associated displays.
Switches which affect the SCS are discussed in the following para-
graphs.
AUTO RCS SELECT Switches.
Power to the RCS control box assembly is controlled by 16 switches
(O-14, figure 3-1). Individual engines may be enabled or disabled
as required. Power to the attitude Control logic is also controlled
in this manner, which thereby controls all attitude hold and/or
maneuvering capability using SCS electronics (automatic coils).
Note
The automatic coils cannot be activated until the RCS
ENABLE is activated either by the MESC or manually.
DIRECT Switches.
Two DIRECT switches (W-35, figure 3-1) provide for manual
Section III Emergency Detection and Procedures
control of the SM RCS engines. Switch 1 controls power to thedirect solenoid switches in rotational controller 1 and switch 2
controls power to the direct solenoid switches in rotational controller
2. In the down position switch 1 receives power from MNA and
switch 2 receives power from MNB. In the up position both switches
receive power from both MNA and MNB. Manual control is
achieved by positioning the rotational control hardover to engage
the direct solenoids for the desired axis change.
ATT SET Switch.
The ATT SET switch (R-40, figure 3-1) selects the source of totalattitude for the ATT SET resolvers as outlined below.
Position Function
UP IMU Applies inertial measurement unit (IMU)
gimbal resolver signal to ATT SET resolvers.
FDAI error needles display differences.
Needles are zeroed by maneuvering SC or
by moving the ATT SET dials.
DOWN GDC Applies GDC resolver signal to ATT SET
resolvers. FDAI error needles display dif-
ferences resolved into body coordinates.
Needles zeroed by moving SC or ATT SETdials. New attitude reference is established
by depressing GDC ALIGN button. This willcause GDC to drive to null the error; hence,
the GDC and ball go to ATT SET dial value.
MANUAL ATTITUDE Switches.
The three MANUAL ATTITUDE switches (T-34, figure 3-1) are
only operative when the SC is in the SCS mode of operation.
Position Description
ACCEL CMD Provides continuous RCS firing as long as therotational controller is out of detent.
RATE CMD Provides proportional rate command from rota-
tional controller with inputs from the BMAG's in
a rate configuration.
MIN IMP Provides minimum impulse capability through therotational controller.
LIMIT CYCLE Switch.
The LIMIT CYCLE switch (T-37, figure 3-1), when placed in the
LIMIT CYCLE position, inserts a psuedo-rate function which
provides the capability of maintaining low SC rates while holding
the SC attitude within the selected deadband limits (limit cycling).
This is accomplished by pulse-width modulation of the switching
amplifier outputs. Instead of driving the SC from limit-to-limit
with high rates by firing the RCS engines all the time, the engines
are fired in spurts proportional in length and repetition rate to
the switching amplifier outputs. Extremely small attitude corrections
could be commanded which would cause the pulse-width of the
resulting output command to be of too short a duration to activate
the RCS solenoids. A one-shot multivibrator is connected in parallel
to ensure a long enough pulse to fire the engines.
RATE and ATT DEADBAND Switch.
The switching amplifier deadband can be interpreted as a rateor an attitude (minimum)deadband. The deadband limits are a
function of the RATE switch (T-39, figure 3-1). An additionaldeadband can be enabled in the attitude control loop with the
AT]- DEADBAND switch (T-38, figure 3-1) See figure 3-3 for
relative rates. The rate commanded by a constant stick deflection
(proportional rate mode only) is a function of the RATE switch
position. The rates commanded at maximum stick deflection (soft
stop) are shown in figure 3-4.
3-13
Section III Emergency Detection and Procedures
RATE ATT DEADBAND
RATE DEADBAND SWITCH POSITIONSWITCH POSITION
LOW
HIGH
+0,2
+ 2.0
MINIMUM MAXIMUM
+ 0.2 ° + 4.2 o
+4.0 ° +8.0 °
CH.20139
Figure 3-3
RATE
SWITCH POSITION
LOW
HIGH
MAXIMUM PROPORTIONAL
RATE COMMAND
PITCH AND YAW
0.7°/SEC
7.0°/SEC
ROLL
0.7O/SEC
20.0O/SEC
Figure 3-4
CH-20140
EM! FUNCTION SWITCH )PERATION
OPERATIONAL
MODE
_V MODE
SELF TEST
AND ENTRYMODE
3-14
SWITCH
SELECTION
START AT _v
AND ROTATE
CLOCKWISE.
START AT NO. 1
AND ROTATE
COUNTERCLOCKWISE
SWITCH
POSITION
_V
Z_V SETVH-_f_'_G
AV TEST
NO. 1
NO. 2
NO. 3
NO. 4
NO. 5
RI NG SET
VO SET
ENTRY
OFF
DESCRIPTION
CORRECT PORTION FOR SPS THRUST MONITORING
( _V DISPLAY).
ENABLES USE OF EMS/AV SET SWITCH TO SLEW_V/RANGE DISPLAY TO INITIAL CONDITION FOR
Z_V TEST AND SPS THRUST MONITORING. PROVIDES
VHF RANGING INFORMATION FOR Z_V//RANGE DISPLAY.
VERIFIES CORRECT OPERATION OF:
1. SPS THRUST LAMP
2. AV DISPLAY (AND COUNTDOWN ELECTRONICS).
(SEE _V SET POSITION ABOVE.)
3, THRUST-OFF COMMAND
TEST EMS FOR DECELERATION <.05G, (NO LAMPSI LLUMI NATED).
DECELERATION >.05G (.0SG LAMP SHOULD ILLUMINATE.)
DECELERATION < .262G.
1. .05G LAMP ILLUMINATES IMMEDIATELY.
2. TEN SECONDS LATER BOTTOM LAMP ON RSI ILLUMI NATED.
3. ENABLES SLEWING OF _V/RANGE DISPLAY.
EMS SYSTEM TEST.
1. AV/RANGE DISPLAY DRIVESTO0 +0.2iN 10SECONDS.
2. VELOCITY SCROLL DRIVES RIGHT TO LEFT.
3. G SCRIBE DRIVES DOWN TO 90 IN 10 SECONDS.
4. .05G LAMP ON.
DECELERATION >.262G.
1. ILLUMINATES .05G LAMP IMMEDIATELY.
2. TEN SECONDS LATERTOP LAMP ON RSt ILLUMINATED.
3. G SCRIBE DRIVES UP TO 0.28 + 0.01G.
4. ENABLES SLEWING SCROLL TO 37,000 FPS.
ENABLES SLEWING /W/RANGE DISPLAY TO INITIAL
CONDITION USING EMS/_.V SET SWITCH. G SCRIBE
DRIVES VERTICALLY TO 0 + 0.1G.
ENABLES SLEWING VELOCITY SCROLL TO INITIAL
COUNTDOWN USING EMS/ _V SET SWITCH.
OPERATIONAL POSITION FOR EMS ENTRY DISPLAY
FUNCTIONS.
DEACTIVATES EMS EXCEPT FOR SPS THRUST ON LIGHT
AND ROLL ATTITUDE INDICATOR.
Figure 3-5
CH-20142
SC CONT Switch.
The SC CONT switch (W-38, figure 3-1) selects the spacecraft
control as listed below:
Position Description
CMC Selects the G&N system computer controlled SC
attitude and TVC through the digital auto-pilot. An
auto-pilot control discrete is also applied to CMC.
SCS The SCS system cot_trols the SC attitude and TVC.
BMAG MODE Switches.
The BMAG MODE switches (Z-32, figure 3-1) select displays for
the FDAI using SCS inputs.
Position Description "
RATE 2 BMAG set no. 2 provides rate damping and the rate
displays on the FDAI.ATT 1 BMAG set no. 1 is uncaged providing attitude hold
and attitude error display on the FDAI while,
RATE 2 Set no. 2 provides rate damping and the rate display.RATE 1 BMAG set no. 1 provides rate damping and the rate
displays on the FDAI.
ENTRY MONITOR SYSTEM (EMS).
The EMS provides displays and controls to show automatic primary
guidence control system (PGNCS) entries and AV maneuvers and
to permit manual entries in the event of a malfunction. There are
five displays and/or indicators which monitor automatic or manual
entries and four switches to be used in conjunction with these
displays.
ENTRY EMS ROLL Switch.
The ENTRY EMS ROLL switch (AE-37, figure 3-1) enables the
EMS roll display for the earth reentry phase of the flight.
ENTRY .05 G Switch.
Illumination of the .05 G light (K-48, figure 3-1) is the cue for
the crew to actuate the .05 G switch (AE-38). During atmospheric
reentry (after .05 G), the SC is maneuvered about the stability
roll axis rather than the body roll axis. Consequently, the yaw
Section III Emergency Detection and Procedures
rate gyro generates an undesirable signal. By coupling a component
of the roll signal into the yaw channel, the undesirable signal is
cancelled. The .05 G switch performs this coupling function.
EMS FUNCTION Switch.
The EMS FUNCTION switch (G-45, figure 3-1) is'a 12 position
mode selector switch, used as outlined in figure 3-5.
EMS MODE Switch.
The EMS MODE switch (I-44, figure 3-1) performs the following
functions in the positions indicated:
NORMAL--Enables EMS accelerometer.
STBY-Inhibits operation in all but AV SET, RNG SET, and Vo
SET positions of FUNCTION switch.
BACKUP VHF RNG
1. A manual backup to automatic .05G trigger circuits that starts
scroll drive and RANGE integrator display drive circuits. Also
backup to TVC MODES for velocity monitoring.
2. Does not permit negative acceleration pulses into countdown
circuits.
3. Enables'VHF ranging information to be displayed on .XV/
RANGE display.
Threshold Indicator (.05 G Light).
The threshold indicator (.05G light) (K-48, figure 3-1) provides
the first visual indication of total acceleration sensed at the reentry
threshold (approximately 290,000 feet). Accelerometer output is
fed to a comparison network and will illuminate the .05 G lamp
when the acceleration reaches .05 G. The light will come on notless than 0.5 sec or more than 1.5 sec after the acceleration reaches
.05 G and turns off when it falls below .02G (skipout).
Corridor Indicators.
By sensing the total acceleration buildup over a given period of
time, the reentry flight path angle can be evaluated. This data
PERIOD MODE DESCRIPTION NOTE
PAD TO 1:01
1:01 TO 100,000 FEET (1:50)
100,000 FEET TO LET JETT
(1:50) (2:51)
LET JETT TO &R = -1134 NM
(2:5 I) (9: 32)
DR = -1134 NM TO _,R = -450 NM
(9:32) (9:49)
_,R = -450 NM TO INSERTION
(9:49) (10:01)
COl CAPABILITY TO INSERTION
(9:47) (10:01)
MODE IA
MODE IB
MODE IC
MODE II
MODE IliA
MODE IIIB
CSM NO GO/$LV LOFTED
MODE IV
CSM GO
LET
LOW ALT
LET
MED ALT
LET
HIGH ALT
FULL-LIFT
FULL-LIRT
SPS POSIGRADE
SPS RETRO
FULL-LIFT
SPS TO ORBIT
(i)
(I)
(i)
(i) (2)
(i) (2)
(i) (2)
(i) (3)
NOTES
(I) EVENT TIMES (MI NUTES: SECONDS) ARE APPROXIMATIONS
(2) _R = CMC SPLASH ERROR WITH HALF-LIFT. _R = -450 NM AND FULL LIFT WILL LAND SC
AT TARGET.
(3) FOR POSITIVE h AND S-IVB CUTOFF BEYOND THE 5 MINUTE TO APOGEE LINE (CREW
CHART) AN APOGEE KICK MANEUVER WOULD BE RECOMMENDED FOR THE MODE IV.
CH-2014
Figure 3-63-15
Section III Emergency Detection and Procedures
is essential to determine whether or not the entry angle is steep
enough to prevent superorbital "skipout." The two corridor indica-tor lights (K-45 and L-45, figure 3-1) are located on the face of
the roll stability indicator (L-45). If the acceleration level is greater
than 0.262 G at the end of a ten second period after threshold
(.05 G light ON), the upper light will be illuminated. It remains
ON until the G-level reaches 2 G's and then goes OFF. The lower
light illuminates if the acceleration is equal to or less than 0.262
G at the end of a ten second period after threshold. This indicates
a shallow entry angle and that the lift vector should be down for
controlled entry, i.e., skipout will occur.
Roll Stability Indicator.
The roll stability indicator (L-45, figure 3-1) provides a visual
indication of the roll attitude of the CM about the stability axis.
Each revolution of the indicator represents 360 degrees of vehiclerotation. The display is capable of continuous rotation in either
direction. The pointer up position (0 degrees) indicates maximum
lift-up vector (positive lift) and pointer down (180 degrees) indicates
maximum lift-down vector (negative lift).
G-V Plotter.
The G-V plotter assembly (H-49, figure 3-1) consists of a scroll
of mylar tape and a G-indicating stylus. The tape is driven from
right to left by pulses which are proportional to the acceleration
along the velocity vector. The stylus which scribes a coating on
the back of the mylar scroll, is driven in the vertical direction in
proportion to the total acceleration. The front surface of the mylar
scroll is imprinted with patterns consisting of "high G-rays" and
"exit rays." The "high-G-rays" must be monitored from initial entry
velocity down to 4000 feet per second. The "'exit rays" are significant
only between the entry velocity and circular orbit velocity and are,
therefore, only displayed on that portion of the pattern. The im-
printed "high-G-rays" and "exit rays" enable detection of primary
guidance failures of the type that would result in either atmospheric
exits at supercircular speeds or excessive load factors at any speed.
The slope of the G-V trace is visually compared with these rays.
If the trace becomes tangent to any of these rays, it indicates a
guidance malfunction and the need for manual takeover.
V/RANGE Display.
The AV/RANGE display provides a readout of inertial flight path
distance in nautical miles to predicted splashdown after .05G. The
predicted range will be obtained from the PGNCS or ground
stations and inserted into the range display during EMS range
set prior to entry. The range display will also indicate AV (ft/sec)
during SPS thrusting.
ABORT MODES AND LIMITS.
The abort modes and limits listed in figures 3-6 and 3-7 are based
on a nominal launch trajectory. The nominal launch phase callouts
are listed in figure 3-8.
Note
More specific times can be obtained from current mis-sion documentation.
3-16
EMERGENCY MODES.
Aborts performed during the ascent phase of the mission will be
performed by using either the Launch Escape System or the Service
Propulsion System.
LAUNCH ESCAPE SYSTEM.
The Launch Escape System (LES) consists of a solid propellant
launch escape (LE) motor used to propel the CM a safe distance
from the launch vehicle, a tower jettison motor, and a canard
subsystem. A complete description on use of the system can be
found in the specific mission Abort Summary Document (ASD).
A brief description is as follows:
Model IA Low Altitude Mode.
In Mode IA, a pitch control (PC) motor is mounted normal to
the LE motor to propel the vehicle downrange to ensure water
landing and escape the "fireball." The CM RCS propellants are
dumped through the aft heat shield during this mode to preclude
damage to the main parachutes. The automatic sequence of majorevents from abort initiation is as follows:
Time Event
00:00 Abort
Ox rapid dumpLE and PC motor fire
00:05 Fuel rapid dump
00:11 Canards deploy00:14 ELS arm
00:14.4 Apex cover jett
00:16 Drogue deploy
00:18 He purge00:28 Main deploy
The automatic sequence can be prevented, interrupted, or replaced
by crew action.
Mode IB Medium Altitude.
Mode IB is essentially the same as Mode IA with the exception
of deleting the rapid propellant dump and PC motor features. The
canard subsystem was designed specifically for this altitude region
to initiate a tumble in the pitch plane. The CM/tower combinationCG is located such that the vehicle will stabilize (oscillations of
-----30 degrees) in the blunt-end-forward (BEF) configuration. Uponclosure of barometric switches, the tower would be jettisoned and
the parachutes automatically deployed. As in Mode IA, crew inter-vention can alter the sequence of events if desired.
Mode IC High Altitude.
During Mode IC the LV is above the atmosphere. Therefore, the
canard subsystem cannot be used to induce a pitch rate to the
vehicle. If the LV is stable at abort, the LET is manually jettisoned
and the CM oriented to the reentry attitude. This method provides
a stable reentry but requires a functioning attitude reference. With
a failed platform the alternate method will be to introduce a five
degree per second pitch rate into the system. The CM/tower combi-nation will then stabilize BEF as in Mode lB. The LES would
likewise deploy the parachutes at the proper altitudes.
Mode II.
The Sm RCS engines are used to propel the CSM away fromthe LV. When the CSM is a safe distance and stable, the CM
isseparatedfromthe SM and maneuvered to a reentry attitude.
A normal entry procedure is followed from there.
SERVICE PROPULSION SYSTEM.
The Service Propulsion System (SPS) aborts utilize the Service
Module SPS engine to maneuver to a planned landing area, or
boost into a contingency orbit. The SPS abort modes are:
Mode IliA.
The SPS engine is used for a posigrade maneuver to fly over the
cold water in the north Atlantic and land at a predetermined point.
The duration of the SPS burn is dependent on the time of abort
initiation. Upon completion of the burn, normal entry procedureswill be followed.
Section III Emergency Detection and Procedures
Mode IIIB.
The SPS engine is used to slow the CSM combination (retrograde
maneuver) so as to land at a predetermined point in the Atlantic
Ocean. The length of the SPS burn is dependent upon the time
of abort initiation. Upon completion of the retro maneuver, normal
entry procedures will be followed.
Mode IV.
The SPS engine can be used to make up for a deficiency in insertion
velocity up to approximately 1300 feet per second. This is ac-
complished by holding the CSM in an inertial attitude and applyingthe needed AV with the SPS to acquire the acceptable orbital
velocity. If there is no communication with MCC, the crew cantake over manual control and maneuver the vehicle using onboard
data.
RATES
1. PITCH AND YAW 5 :_0.S DEG/SEC
UFTOFF (1"+0) TO T + I MIN 40 SEC
T + 1 MIN 40 SEC TO S*IVB CUTOFF 10 _0.S DEG/SEC
2. ROLL 20:1:0.5 DEG/SEC
LIFTOFF TO S-IVB CUTOFF
PLATFORM FAILURE
1. DURING S-IB POWERED FLIGHT THE TWO CUES FOR PLATFORM FAILURE REQUIRING AN
IMMEDIATE SWITCHOVER ARE:
a. LV GUID LT - ON
b. LV RATE LT - ON
2. AFTER LV RATE SWITCH DEACTIVATION THE PRIMARY CUE IS:
LV GULD LT - ON
THE SECONDARY CUES ARE:
a. FDAI ATTITUDE
b. LV RATES
c. GROUND CONFIRMATION
AUTOMATIC ABORT LIMITS (LIFTOFF UNTIL DEACTIVATION AT T + 1 MIN 40 SEC)
1. RATE PITCH - YAW S :LO.S DEG/SEC
ROLL 20 :L0.5 DEG/SEC
2, ANY TWO ENGINES FAIL
3. CM TO IU BREAKUP
S-IB ENGINE FAILURE (SUBSEOUENT TO AUTO ABORT DEACTIVATE AT T + 1 MIN 40 SEC0
1. SINGLE ENGINE FAILURE
2. SIMULTANEOUS LOSS TWO OR MORE ENGINES
S-IVB ENGINE FAILURE
S-IVB DIFFERENTIAL TANK PRESSURE LIMITS
_P (ORBITAL COAST) LH 2 > LO 2 = 26 psld
LO 2 > LH 2 = 36 psid
LO 2 > 50 psla
CONTINUE MISSION.
ABORT F LV CONTROL IS LOST.
ABORT (MODE IC, II, Ill, OR IV)
Figure 3-7
3-17
Section III Emergency Detection and Procedures
NOMINAL LAUNCH PHASE VOICE CALLOUT! BOOST TO ORBIT
PROG TIMEr
-00:03
+00:01
0O: 10
0O: t2
0O:30
01:01
OO:50
0O: 55
01:40
01:50
02:10
02:19
02:22
02:23
02:24
02:27
02:51
02:52
04:0O
05:00
06:0O
07: 0O
08:00
09:00
09:32
09:47
09:49
10:01
STA
LCC
LCC
CDR
LCC
CDR
CDR
MCC
CMP
CDR
CMP
CMP
MCC
MCC
CDR
CDR
CDR
CDR
CDR
CMP
MCC
CMP
CDR
CDR
CDR
MCC
CDR
CDR
CDR
CDR
MCC
CDR
MCC
MCC
MCC
CDR
MCC
ACTION/ENTRY *REPORT
IGNITION*
LI_OFF"
CLOCK START*
CLEAR TOWER*
ROLL & PITCH START*
ROLL COMPLETE*
MODE IB*
PRPLNT DUMP-RCS CMD
MONITOR _ TO T+1:40
MONITOR CAB1N PRESSURE DECREASING
EDS AUTO-OFF*
EDS ENG-OFF
EDS RATES-OFF
KEY VB2E, N62E
MODE 1C* (BASED ON 100,000 FT)
GO/NO GO FOR STAGING +
GO/NO GO FOR STAGING*
INBOARD OFF*
OUTBOARD OFF*
S-IB/S-IVB STAGING*
S-IVB IGNITION COMMAND
S-IVB 65%
TWR JETT (2)-ON* (IFTFF >I+20)
MODE lJ*
c_/PC-PC
MAN ATT (PITCH)-RATE CMD
GUIDANCE INITIATE*
REPORT STATUS"
TRAJECTORY STATUS*
REPORT STATUS*
REPORT STATUS*
REPORT STATUS +
GMBL MOT (4)-START-ON
CHECK GPI (/V_bMENTARILY)
REPORT STATUS*
GO/NO GO FOR STAGING*
GO/NO GO FOR STAGING
MODE Ilia
MODE IV
MODE IIIB
SECO*
rNSURE ORBIT
KEY RLSE TO N44
INSERTION*
OPTION/EVENT
UMBILICAL DISCONNECT
CMC TO Pll
DET & MET START
ABOVE LAUNCH TOWER
ROLL AND PITCH PROGRAM START
ROLL PROGRAM COMPLETE
IF NO DECREASE BY 17,000 FEET
DUMP MANUALLY
NO AUTO ABORT LIGHT-ON
SYSTEMS STATUS
ENGINE LIGHTS NO. 5, 6, 7, & 8 ON
LIFTOEF LIGHT-OUT
NO AUTO ABORT LIGHT-OUT
ENG LIGHTS I THRU 4-ON
ENG LIGHTS OUT
ENG LIGHT NO. 1 ON
ENG LIGHT NO. t OUT
TOWER JETTISON
IGM START
I NSURE ANGLES CORRECT
SYSTEMS STATUS
ENG LIGHT NO. I ON MOMENTARILY
tEVENT TIMES ARE APPROXIMATIONSCH-20144
Figure 3-83-18
SECTION IV ==_=='
S-IB STAGE
TABLE OF CONTENTS
Introduction ........................................................................ 4-1
Structure; ............................................................................ 4-1
Propulsion .......................................................................... 4-7
Control Pressure System .................................................. 4-33
S-IB Hydraulic System ....................................................... 4-37Electrical ............................................................................ 4-43
Instrumentation .................................................................. 4-52
Environmental Conditioning .............................................. 4-56Ordnance ............................................................................ 4-59 ;-IB ;TAGE
INTRODUCTION.
The function of the S-IB stage is to boost the upper stages and
spacecraft through a predetermined trajectory that will place them
at the proper altitude and attitude, with the proper velocity, at
S-IVB stage ignition. The major S-IB stage assemblies, figures 4-1
through 4-3, are the tail unit with eight fins and eight H-I engines,
the nine propellant tanks, the second stage adapter (spider beam
unit), and associated mechanical and electrical hardware discussed
under specific systems in this section. For a summary of S-IB stage
data, see figure 4-4.
STRUCTURE.
Figure 4-5 shows the primary, load-carrying structural subassemb-
lies of the S-IB stage combined with its tail unit heat and flame
shields, engine flame curtains, lox and fuel tank firewalls, andsecond stage adapter seal plate. Separate figures show the unique
design details of the tail unit heat shield, and also the spider beamand lox fitting reinforcements employed as a result of qualification
testing. The stage structure was designed to provide a safety factorof 1.10 on yield and 1.40 on ultimate, with a dry stage weight
of 85,745 Ibm. The adequacy of the 1.40 safety factor (ultimate)
has been demonstrated by all load-carrying structural subassem-
blies. From a reliability point of view, the stage structure is a simple,
passive system, and its reliability prediction is based solely on
whether or not strength will exceed load. The 1.40 ultimate safety
factor, a conservative compilation of two-sigma and three-sigma
loads and allowables, plus the complete analysis and test program
demonstrates a reliability assessment several times greater than
the reliabilities of the other stage systems. Thus, with respect to
the rest of the onboard systems, the structure has been assumed
to be 100 percent reliable within the performance limits established
by the CEI specification, and is considered so for purposes of
calculating total stage reliability. The principal functional require-
ment of the stage structure is to provide adequate tankage and
framework to support the other flight systems and to provide
adequate support of the upper stages, both on the pad and in
flight. The evolutionary structural changes resulting from design
analyses, test results, static firings, and flight performance data
are summarized separately in the discussion pertaining to the indi-vidual structural elements.
PROPELLANT TANKS.
The nine propellant tanks that cluster to form the main body of
the stage are modifications of proven designs from the Redstone
and Jupiter vehicles and have performed successfully on all Saturn
I and IB flights. The individual tanks are constructed of cylindrical
sections built up of mechanically-milled, butt welded, aluminum
alloy skin segments that are internally reinforced with rings to form
a monocoque type of construction. The material used to constructthe tanks is the now readily weldable 5456 aluminum alloy in the
H343 temper. The use of this alloy, made possible by welding
ANTISLOSH
BAFFLES
tALL TANKS)_
FUEL FILL &
(8 PLA IIES_LA_
LOX SUCTION LINE
C""
PLATE
ANTENNA
(2 PLACES)
(8 PLACES)
DUCT
(FUEL TANKS 1,2,&3)
(4 PLACES)
)UTER LOX TANK
(4 PLACES)
LOX TANK
OUTRIGGER
(8PLACES)
LICAL
SHROUD
(8PLACES)
Figure 4-1
4-1
Section IV S-IB Stage
S-IB STAGEFORWARDDETAIL
MANIFOLD
INSTRUMENT
COMPARTME NT
ENVIRONMENTAL
CONDITIONING DUCTS _
VALVE (4 PLACES)
\
iiJ
TELEMETRY_ANTENNA'/(2 PLACES)--"
COMPARTMENT
(2 PLACES)
LOX VENT AND
.OX
L-
NK
FUEL TANK
F-2
LOX TANK
L-3
Hetes
REDUNDANT ANTENNAS
LOCATED AT POS I AND Ill
_> FUEL TANKS F-3 AND F-4
[_ FORWARD SKIRTS OF FUEL TANKS F-I AND F-2
_ TELEMETRY ANTENNA LOCATED AT POS |1 AND IV
I I
FUEL TANK
F-3
_ANTISLOSH
BAFFLES
LOX TANK
L-4
HELIUM
STORAGE SPHERE (2 PLACES)
VENT
VALVE(2 PLACES)
RANGE
SAFETY COMMANDSYSTEM ANTENNA _]
(4 PLACES)
C.H 14206-1
4-2 Figure 4-2
Section IV S-IB Stage
S-IB STAGEAFTDETAIL
FUEL FILL AND
DRAIN
FUEL FILL AND
FUEL
TAN
F-I
LOX TANK
L-2
FUEL
PREVALVES-_
CENTER LOX
TANK EUMP --
LOX INTERCONNECT
L
LOX
FUEL TANK
F-2 LOX TANK
L-3
I FUEL SENSOR
DEPLETION
LEVEL_NSO.<:_
FUEL TRANSFER
ASSEMBLY
FUEL
TANK
F-3
©
'_LOX FILL ANDDRAIN NOZZLE
_LOX FILL ANDDRAIN VALVE
LINES
CONDITIONING MANIFOLD
Moles
_:> IDENTICAL INSTALLATIONIN FUEL TANK F-4
[_ IDENTICAL INSTALLATION
IN FUEL TANK F-4 AND
LOX TANKS L-2 AND L-4
GH 14209
Figure 4-34-3
SectionIVS-IBStage
S-IB STAGE DATA SUMMARY
DIMENSIONS
LENGTH
DIAMETER
AT PROPELLANT TANKS
AT TAIL UNIT ASSEMBLY
AT FINS
FIN AREA
MASS
DRY STAGE
LOADED STAGE
AT SEPARATION
ENGINES, DRY, LESS
INSTRUMENTATION
INBOARD, PLUS TURNBUCKLES
OUTBOARD, LESS HYDRAULICS
PROPELLANT LOAD
ENGINES
BURN TIME
TOTAL THRUST (SEA LEVEL)
PROPELLANTS
MIXTURE RATIO
EXPANSION RATIO
CHAMBER PRESSURE
OXIDIZER NPSH (MINIMUM)
FUEL NPSH (MINIMUM)
GAS TURBINE PROPELLANTS
TURBOPUMP SPEED
ENGINE MOUNTING
INBOARD
OUTBOARD
80.2 FT
21.4 FT
22.8 FT
40.7 FT
53.3 FT 2 EACH OF 8 FINS
84,521 LBm
997,127 LBm
95,159 LBm
2,003LB m EACH
1980 LBm EACH
912,606 LBm (408,000 KG)
141 SEC (APPROX)
1.64 MLBf
LOX AND RP-1
2.23:1 + 2%
B: 1
702 psia
35 FT OF LOX OR 65 psia
35 FT OF RP-1 OR 57 psla
LOX AND RP-I
6680 RMP
32 IN. RADIUS, 3 DEG
CANT ANGLE
95 IN. RADIUS 6 DEG
CANT ANGLE
[ ATA SUMMARYHYDRAULIC SYSTEM
ACTUATORS (OUTBOARD ONLY)
GIMBAL ANGLE
GIMBAL RATE
GIMBAL ACCELERATION
PRESSURIZATION SYSTEM
OXIDIZER CONTAINER
FUEL CONTAINER
OXIDIZER PRESSURE
PREFLIGHT
INFLIGHT
FUEL PRESSURE
PREFLIGHT
INFLIGHT
ULLAGE
OXIDIZER
FUEL
ENVIRONMENTAL CONTROL SYSTEM
PREFLIGHT AIR CONDITIONING
PREFLIGHT GN 2 PURGE
ASTRIONICS SYSTEMS
GUIDANCE
TELEMETRY LINKS
TRACKING
ELECTRICAL
RANGE SAFETY SYSTEM
2 PER ENGINE
+ 8 DEG SQUARE PATTERN
15 DEG/SEC IN EACH PLANE
1776 DEG/SEC 2
INITIAL HELIUM FROM GROUND SOURCE;
S-IB BURN, GOX
HELIUM
58 psla
50 psia
17 psig
15 TO 17 psig
1.5%
2.0%
AFT COMPARTMENT & INSTRUMENT
COMPARTMENTS FI & F2
AFT COMPARTMENT & INSTRUMENT
COMPARTMENTS F1 & F2
PITCH, ROLL, AND YAW PROGRAM THRU
THE IU DURING S-IR BURN
FM/FM, 240.2 MHz; PCM/FM, 256.2 MHz
ODOP
BATTERIES, 28 Vdc (2 ZINC-SILVER OXIDE);
MASTER MEASURING
VOLTAGE SUPPLY, 28 Vdc TO 5 V&:.
PARALLEL ELECTRONICS, REDUNDANT
ORDNANCE CONNECTIONS.
Notes ALL MASSES ARE APPROXIMATE.
Figure
advancements, allows for considerable tank weight reduction over
the 5086 and 5052 alloys used to construct the Jupiter and Redstone
tanks, respectively. Tank wall thickness varies from top to bottom
in relation to stress distributions. Hemispherical bulkheads are
welded to the forward and aft end of the cylindrical sections, and
a sump is welded to the aft bulkhead. A pressurization and ventmanifold is fastened to the forward bulkhead of each of the lox
tanks. A cylindrical skirt reinforced with longerons is attached to
the forward and aft bulkheads to complete a basic tank. The eightouter tanks are 70 in. in diameter and contain lox and fuel alter-
nately. The center tank is 105 in. in diameter and contains lox. The
center lox tank is bolted to the spider beam and is attached tothe tail barrel with huck bolts. Ball and socket fittings attach the
aft ends of the 70-in. lox tanks to the tail unit. Banjo fittings and
studs rigidly secure the forward ends of the lox tank to the spider
beam. The fuel tanks are supported by ball and socket fittings
at the tail unit. During shipment, banjo fittings rigidly secure the
fuel tanks to the tail unit; however, they are removed before flight.The forward ends of the fuel tanks are mounted to the spider beam
unit by sliding pin connections, allowing the lox tanks to shorten
CH-14484-2
4-4
due to thermal contraction when loaded. In summary, the major
structural improvements incorporated into the propellant tanks are
revised skin gages and reduced bulkhead and frame gages to agree
more closely with stress levels, inversion of the aft dome manholdcover in the center tank, and the addition of GOX interconnect
domes (with the related forward skirt cutouts) and a GOX pres-
surant diffuser. Also, the fuel tanks are painted white instead of
black on SA-206 and subsequent vehicles for thermal reasons.
TAIL UNIT.
Primarily, the tail unit rigidly supports the aft ends of the propellanttank cluster and the vehicle on the launcher; mounts the eight
engines and fins; and provides the thrust structure between the
engine thrust pads and the propellant tanks. Other functions of
the tail unit are to support the lower shroud panels, lox and fuel
bay firewalls, heat shield support beam and panel assemblies,
engine flame curtains, and the engine flame shield support installa-
tion. Unlike the propellant tank units, the tail unit is constructed
with higher strength aluminum alloys of the 7000 series that areheat-treated to the T6 or the T-73 condition.
4-4
Section IV S-IB Stage
FORWARD
SKIRT--_
PRESSURANT
DIFFI
ANTISLOSH
LEADING !
EDGE
PLATE
SUMF
SECOND STAGE ADAPTER
(SPIDER BEAM)
105-INCH CENTER LOX TANK
FRONT
SHEAR ANGLE
REARSPAR_
NTERNAL
REINFORCING
RING
HEMISPHERICAL
BULKHEAD
TRANSFER
MANIFOLD
INTERCONNECT
FLANGE
AFT SKIRT
TUBE (2)
)LDDOWN
FITTING
UPPER SHROUD
PANEL
UPPER RING
SEGMENT
LOWER RING
SEGMENT
HONEYCOMB
FiREWALL
LOWER SHROUDPANEL ASSEMBLY (8)
(SMOOTH SKINNER
INBOARD ENGINE
SHEAR WEB ASSEMBLY
INBOARD ENGINE
THRUST PAD
TORQUE BOX HEAT
SHIELD PANEL (16)
OUTBOARD ENGINE
FLAME CURTAIN
SEAL RING
Figure 4-5
CABLE ROUTING
DUCT (FUEL TANKS
1,2,&3 ONLY)
SUMP-_
AFT SKIRT_.._
REINFORCING RINGS _
BARREL ASSEMBLY
SKIN I
-....=
7D-INCH OUTER TANK
(TYPICAL)
TION ASSEMBLY
OUTRIGGER
ASSEMBLY (4)
RESTRAINING
BRACKETTRAILING
EDGE
SHIELD
FIN STRUCTURE
(SKIN OMITTED)
_ANTISLOSH
BAFFLES
_SUCTION
LINE
FLANGE
_-AFT
ATTACHING
HOLES
OUTBOARD ENGINE
THRUST PAD (4)
THRUST RING
UPPER THRUST RING
SUPPORT
INSTALLATION
TAIL UNIT STRUCTURE
_ SUPPORTINGBEAM
_i STRUCTURE
SHROUD
PANEL
ASSEMBLY(CORRUGATED)
(8)
fiELD
PANELS (40)INBOARD ENGINE
FLAME CURTAIN (4)
FLAME SHIELD
CH-14225-I
4-5
Section IV S-IB Stage
The 7000 series aluminum alloys used in the tail unit assembly
are not recommended for welded applications because of low weld-
ing efficiencies. These alloys are used in t:npressurized areas wheret':e asser':tb!ing is done with mechanical fastensrs. The tail unit
thrust str_eture configuration lends itself to this definition; and
high strength, heat-treatable 7000 series aluminum alloy forgings,
extrusions, plates and sheets are fabricated into components thatare joined with mechanical fasteners to construct the tail unit
assembly.
Because of the susceptibility of the high strength aluminum alloys
to stress corrosion cracking, methods have been employed in the
design and manufacture of the tail unit assembly which minimize
the danger of failure due to stress corrosion cracking. Methodsemployed are: heat treatment to the T-73 condition, heat treatment
after heavy machining operations, the use of closed die forgings,
and the use of adequate final protective finishes.
The tail unit consists of a barrel assembly, 105 in. in diameter,
that directly supports the center propellant tank, encloses the in-board engine thrust beams, and acts as the hub for the four thrust
support outriggers and the four fin support outriggers. The four
thrust support outriggers also act as fin support outriggers. The
fin support outriggers are similar to the thrust support outriggers
but differ mainly in that they have no thrust support beam or
actuator support beam. The outer ends of the outriggers are spanned
by upper and lower ring segments and eight upper shroud panels
to form the basic thrust structure. Eight smooth and eight cor-
rugated lower shroud panels are attached to the aft end of the
thrust structure to form a compartment for the eight H-I engines.Lox and fuel bay firewall panels are installed to cover the space
between the outrigger assemblies and the space over the aft endof the barrel assembly.
A reinforcing beam structure is fitted into the aft end of the lower
engine shroud assembly. The heat shield panels, engine flame
curtains, and flame shield support installation are attached to the
beam structure. Figure 4-6 shows the unique design details of the
tail unit heat shield. Unlike most other honeycomb composites used
in the vehicle utilizing phenolic cores that are adhesively bonded
to face sheets and are limited by the upper and lower temperature
constraints on the adhesive system, the heat shield honeycombcomposite consists of both corrosion-resistant steel foil cores and
thin face sheets that are joined by a brazing process. The 0.25-in.square-cell core is brazed to both the inner and outer face sheets.
has a layer thickness of 1.00 in., and acts as the chief structural
core member of the composite. The 0.50-in. square-cell core is
brazed to only the outer face sheet, has a layer thickness of 0:25-in.,
and acts as the thermal insulation retaining member of the compos-
ite structure. M-31 insulation is trowled into the retaining core
cells. Laboratory tests have generally demonstrated that. compared
with adhesively bonded honeycomb composites, brazed honeycomb
composites are over 100 percent greater in tensile strength, over
75 percent greater in core shear strength, over 20 percent greater
in edgewise compression strength, and equal in flatwise compression
strength. This heat shield design provides a lighter panel with
increased stiffness which greatly improves the retention of the M-31
insulation material. Successful results of laboratory testing and static
tests of the S-I-10 and S-IB-3 through S-IB-7 stages have fullyqualified this heat shield panel design. In addition to the new
configuration heat shield assembly, the tail unit incorporates new
fin attachment fittings, gage reduction of sheet-metal and framing,
and removal of the engine skirts from the lower shroud assembly.
FINS.
The eight nl,_, ,_" _emi-monocoque construction, provide aerody-
namic stability ih id-region of first stage flight and support
the vehicle on the ch pad prior to ignition and during the
hold-down period _. ignition. A fin is fastened mechanically
HEAT SHIELD DETAILS
4-6
to each of the four thrust support outriggers and the four fin support
outriggers. A heat shield is attached to the trailing edge to protect
the fin from engine exhaust, and a plate is fastened to the tip
of the fin between the leading edge and the heat shield. Skin panels
are riveted to the ribs and spars, completing the structure andforming a smooth aerodynamic surface. The fins used on the S-IB
stage are identical, and of a completely new configuration, replacing
the arrangement of four large-fins and four stub-fins used on the
S-I stages.
SPIDER BEAM UNIT.
The spider beam unit holds the propellant tank cluster together
at the forward end and attaches the S-IB stage to the S-IVB aft
interstage. The five lox tank units are rigidly attached to the spider
beam while the fuel tank units are attached with sliding pin connec-
tions. Structurally, the spider beam consists of a hub assembly,
to which eight radial beams are joined with upper and lower splice
plates by mechanical fasteners. The outer ends of the radial beams
are spanned by crossbeams and joined with upper and lower splice
plates by mechanical fasteners. Like the tail unit thrust structure
the spider beam is constructed of extrusions and fittings made ofhigh strength, heat-treatable, aluminum alloys of the 7000 seriesthat are heat-treated to the T6 condition. To form an aft closure
for the S-IVB stage engine compartment, twenty-four honeycomb
composite seal plate segments of approximately 0.05-in. thickness
are fastened to the forward side of the spider beam. The seal plate
honeycomb composite consists of 5052 aluminum alloy fail core
material adhesively bonded to 7075-6 aluminum alloy face sheets
to form thinner and lighter panels than those used on the S-I stages.During qualification testing of the S-IB stage spider beam, a failure
of the lox tank fitting occurred. Figure 4-7 shows the radial beam
reinforcing angle and bracket, crossbeam web stiffening brackets,
and reinforced mounting stud flange incorporated to fix each of
the eight lox tank fittings. Other changes incorporated into the
spider beam design for the S-IB stages are the reduction of beam
gages and the removal of retrorockets, 45-deg fairing, and radial
beam tips. This unit is qualified and has performed successfully
on all Saturn IB flights.
HEAT SHIELD
// SUPPORT CHANNEL
( " HEAT REFLECTING (_
CRES CELLS, 0.0015 FOIL TAPE LAYERS _ r-_
THICKNESS) "_ _,,,_ _
\ /''-INNER FACE _j_jl
_AZ,NGEO,L\ /SHEETC00,0CRESl1_1LAYERS _. / Ildql _ ASBESTOS
PANEL EDGE TAPE'O002AG'CU'L" 2X_L_ I I/LAYER(0.125
0.25 ,IB _ _i_ii!_i_l!i;il_:.i:.! :_lE!i!_!_ilFi!!!i_i!_i_:.t_i_11r --II_ii_
OUTERPACEV// [cREs. S,EET(o.olo CRES)---/ / / 8OLT---- \
,_j / _ SPACERBRAZINGPOlLLAYER / (0.50D_A10.0015AG-CU-L0 / 0.06 WALL
/ CRES TUBE)M-s_ INSULATrON /IMPREGNATED INTO/HONEYCOMB CORE --.J
INSULATION RETAINING i
HONEYCOMB CORE LAYER /
(0.50-INCH SQUARE CRES /CELLS, 0.0015 FOIL THICKNESS) --J
CH.14344.1
Figure 4-6
;PIDERBEAMFIXES
RADIAL BEAM
REINFORCING
ANGLE
EAM
ASSEMBLY
Figure 4-7
PROPULSION.
The S-IB stage propulsion system consists of an eight-engine cluster
of H-1 engines that burn lox and RP-I fuel to propel the Saturn
IB vehicle during the first boost phase of powered flight. Propellantfrom the lox and fuel tanks feed the H-I engines under tank pressure
to assure the NPSH necessary for satisfactory engine operation.
Boosters S-IB-1 through S-IB-5 used engines developing 200,000
lbf of thrust for a total stage thrust of 1,600,000 lbf. Boosters S-IB-6
and subsequent will use engines developing 205,000 lbf of thrust
for a total stage thrust of 1,640,000 lbf. Four inboard engines are
mounted 90 deg apart (at vehicle positions 1, II, II1, and IV) on
a 32-ini radius from the vehicle longitudinal axis and are canted
3 deg outboard from the vehicle centerline. Four outboard engines
are gimbal mounted 90 deg apart (at fin lines 2, 4, 6, and 8) ona 95-in. radius fromthe vehicle longitudinal axis. The engines cant
outboard 6 deg from the vehicle centerline. Each of the eight engines
is attached by/a gimbal assembly to its thrust pad on the tail unitthrust structure. Inboard engine thrust pads are on the barrel
assembly a:nd outboard engine thrust pads are on the thrust support
outriggci's. Although the inboard engines do not gimbal for vehiclecontrol the gimbal assemblies permit alignment of the engines tothe thrust-structure; two turnbuckles used on each inboard engine,
with the gimbal assembly, align and secure the engine in place.
Two hydraulic actuators and a gimbal assembly secure each out-
Section IV S-IB Stage
board engine to the thrust structure. The actuators attach to an
actuator support beam, which is part of the thrust support outrigger.
The actuators, one mounted in the pitch plane and one in the
yaw plane, gimbal the engine for vehicle attitude control. The
engine gimbal centerline for both outboard and inboard engines
lies in a plane perpendicular to the vehicle longitudinal axis at
vehicle station 100 (figure 4-8). Canting the engines provides stabi-
lity by directing the thrust vectors to common points on the vehicle
longitudinal axis. The outboard engine thrust vectors intersect the
longitudinal axis at vehicle station 1004, while the inboard engine
thrust vectors intersect the longitudinal axis at vehicle station 711.
The difference in cant angles and radii from vehicle centerline
account for the two different intersect points. Directing the thrust
vectors to the vehicle longitudinal axis reduces the possibility of
excessive loading of the vehicle structure in the event of engine(s)
failure during flight.
H-I ENGINE.
The H-1 engine is a single-start, fixed-thrust, bipropellant rocket
engine that burns RP-1 (MSFC-SPEC-342A) fuel and lox (MSFC-
SPEC-399). Calibrated orifices installed in the high-pressure fuel,and at the lox and fuel inlets to the gas generator control valve,
fix the propellant flowrates, which effect the fixed thrust. The engine
STA
I0O4
STA
711
STA
100
Note
ENGINE ANGLES EXAGGERATED
TO SHOW CANT ANGLES AND
THRUST VECTOR INTERSECT POINTS.
OUTBOARD ENGINES INSTALLED
ON 95 IN. RADIUS FROM LONGI-
TUDINAL AXIS; INBOARD ENGINESINSTALLED ON 32 IN. RADIUS FROM
LONGITUDINAL AXIS,
LONGITUDINAL
AXIS
THRUST VECTOR
INTERSECTION
STA 1004
tYP)
THRUST
VECTOR INTERSECTION
STA 711
ENGINE
NO,
ENGINE iiiii_Eii
ENGINE
NO. II NO. 6
ENGINE
NO. 8
IV
POINT
TYP
NO. 3
iI!ili_::"_'-'ENGINE GIMBAL
_ENGINE PLANE STA 100NO. 7
NGINE
NO.2
C-H 14267-1
Figure 4-8
4-7
Section IV S-IB Stage
H-1 ENGINE
93
/94
3O
37
35 34 36
8_8114
10 \
79 80 \
16
73
9
74 54 96 55
90 91 2925 8,
X i.
6 6u;' 51 3_86_i
47
95 57
41
_84
"_' I// so
_ 8_
58
9O
_053 I 2 3 95 7 6 5 11 12 13
24
26
27
15
16
18
20
70 _
68
67
54 55 56 57 58 51 5059
53
66
9064
29
C-H 14231.1
4-8 Figure 4-9 (Sheet 1 of 2)
LEGEND FOR H-1 ENGINEI MAIN FUEL VALVE
2 IGNITION MONITOR VALVE
3 IGNITER FUEL VALVE
4 LPGG CONTROL VALVE
5 CONAX VALVE
6 CONTROL-PRESS FUEL LiNE
7 CONTROL-PRESS FUEL MANIFOLD (CONAX iNLET)8 HIGH-PRESS. FUEL DUCT
9 MLV CLOSING LINE
10 TOPS (3)
11 TOPS SENSING LINE
12 LPGG INJECTOR
13 MEAS T12 (TURBINE RPM)
14 TURBINE EXHAUST HOOD
15 FUEL PUMP VOLUTE
16 LOX PUMP SEAL PURGE PORT
17 GEARBOX PRESSURIZATION CHECK VALVE AND PORT
18 FUEL PUMP INLET
19 MEAS D20 (LOWER GEARCASE LUBE PRESS. SENSING PORT)
20 MEAS D14 (TURBINE INLET PRESS SENSING PORT)21 AFT TURBOPUMP MOUNT SUPPORT
22 MEAS C9 (LPGG COMBUSTION CHAMBER TEMP)
23 SQUIBLESS IGNITERS
24 MEAS C9 ZONE BOX
25 HEAT EXCHANGER GOX OUTLET
26 LPGG COMBUSTION CHAMBER
27 SOLID PROPELLANT GAS GENERATOR
28 SPGG INITIATORS
29 HEAT EXCHANGER
30 THRUST CHAMBER
31 STIFFENING BAND AND TENSION RING
32 ASPIRATOR (OUTBOARD ENGINE ONLY)
33 LUBE DRAIN
34 LOX DRAIN LINE EXTENSION35 LOX SEAL CAVITY DRAIN (SECONDARY)
36 IGNITION MONITOR VALVE DRAIN
37 FUEL DRAIN MANIFOLD DRAIN
38 HEAT EXCHANGE INLET MANIFOLD ASSY
39 FUEL JACKET FILL Q-D COUPLING
40 MEAS D20 (LOWER GEARCASE LUBE PRESS. TRANSDUCER)
41 FUEL ACTUATOR ATTACH POINT (OUTBOARD ONLY)
42 MEAS DI4 (TURBINE iNLET PRESS. TRANSDUCER)
43 FUEL ACTUATOR OUTRIGGER ASSY (OUTBOARD ONLY)
44 MEAS XD35 (GG LOX INJECTOR PRESS. TRANSDUCER)
45 TRANSDUCER PANEL
46 FUEL BOOTSTRAp LINE
47 IMV INLET LINE
48 GG AND IMV CONTROL LINE
49 HYPERGOL CONTAINER
50 LOX BOOTSTRAP LINE
51 THRUST CHAMBER FUEL INJECTOR PURGE LINE
52 TURNBUCKLE ASSY (INBOARD ENGINES ONLY)
53 IGNITER FUEL LINE
54 TURBOPUMP FORWARD MOUNT ASSY
55 HiGH-PRESS. LOX DUCT
56 ACCESSORY DRIVE ADAPTER
57 TC FUEL iNJECTOR PURGE CHECK VALVES (3)
58 GIMBAL ASSEMBLY
59 LOX DOME
60 MEAS D1 (COMBUSTION CHAMBER PRESS. TRANSDUCER)
61 FUEL INLET MANIFOLD
62 LOX ACTUATOR OUTRIGGER ASSY (OUTBOARD ONLY)
63 LOX ACTUATOR ATTACH POINT
64 MEAS XC89 (GEARCASE LUBE TEMP THERMOCOUPLE)
65 FABU FILL Q-D COUPLING
66 FUEL DRAIN Q-D COUPLING
67 FABU DISCHARGE LINE
68 MEAS XC89 ZONE BOX
69 FUEL ADDITIVE BLENDER UNIT
70 FABU FULL INDICATOR
71 TURBOPUMP LUBE FILTER
72 LUBE DRAIN MANIFOLD
73 LOX PUMP VOLUTE
74 LOX PUMP INLET
75 HEATER COVER (LOX BEARING NO. I)
76 MEAS D13 (LOX PUMP INLET PRESS.TRANSDUCER OUTBOARD ONLY)
77 MEAS D12 (FUEL PUMP INLET PRESS. TRANSDUCER OUTBOARD ONLY)
78 MEAS DI2 SENSING LINE (PORT LOCATED ON SUCTION LINE)
79 MEAS D13 SENSING LINE (PORT LOCATED ON SUCTION LINE)
80 TURBINE81 MLV OPENING LINE
82 CUSTOMER CONNECT PURGE PANEL
83 GG LOX INJECTOR PURGE CHECK VALVE
84 HYPERGOL INSTALLED SWITCH ASSY
85 HYPERGOL CARTRIDGE LOCK PiN
86 FUEL JACKET DRAIN PLUG (3)
87 MAIN LOX VALVE
88 LUBE FUEL INLET LINE
89 AUXILIARY DRIVE FAD
90 HEAT SHIELD ASSY
91 GG FUEL INJECTOR PRESS. SENSING PORT
92 MEAS D34 (GG FUEL INJECTOR PRESS. TRANSDUCER)
93 EXHAUST DUCT (INBOARD ENGINES ONLY)
94 FUEL RETURN MANIFOLD (INBOARD AND OUTBOARD ENGINES)
95 CONTROL-PRESS. FUEL MANIFOLD (IFV INLET)
96 MEAS C1 ZONE BOX
97 MEAS C1 (LOX PUMP BEARING NO. 1 TEMP THERMOCOUPLE) C-tl 14272.1
Section IV S-IB Stage
has a regeneratively-cooled thrust chamber with propellant feed
and control components clustered around the forward end of the
combustion chamber (see figure 4-9). A turbopump, driven by a
gas turbine through a gear train, delivers fuel and lox under high
pressure and at high flow rates to the combustion chamber. During
operation, engine control is a function of fuel turbopump dischargepressure. This control-pressure fuel is manifolded from the high-
pressure fuel duct upstream of the main fuel valve (MFV) to the
propellant feed valve actuators. A solid propellant gas generator
(SPGG) spins the turbine to start the H-1 engine and a liquid
propellant gas generator (LPGG), using bootstrap fuel and lox,
sustains engine operation by supplying large volumes of gas to
the turbine. Each engine has its own turbine exhaust system for
expended turbine gases. A heat exchanger in the turbine exhaust
system converts lox to GOX for S-IB stage lox tank inflight pres-surization.
Characteristics.
Inboard engines designated H-1C and outboard engines designated
H-1D have basically the same physical characteristics, except in-
board engines use a partial aspirator, or exhaust duct, for exhausting
turbine gases, while the outboard engines use a peripheral aspirator
on the engine thrust chamber to control the turbine exhaust gas
flow. The outboard engines have outriggers mounted on the thrust-
chamber for hydraulic actuator attachment, while inboard engines
use struts (turnbuckles) attached to thrust-chamber stabilizing lugsfor installation on the tail unit thrust structure. The H-IC and
H-ID engines weigh 1942 and 2003 Ibm (dry) respectively. At start,
engine fluids increase the weight 220 Ibm per engine. Approxi-
mately 166 Ibm of fluids remain in each engine at cut-off. The
engines measure 101.61 in. in length from the bottom of the thrust
chamber to the gimbal assembly mounting face. The outboard
engine aspirator extends an additional 1 in. below the thrustchamber.
H-I Engine History.
H-I engines have flown on all fifteen Saturn I&IB launch vehicles.
The H-I engine has received modifications during the Saturn pro-
gram to uprate thrust and performance capabilities by the addition
of injector baffles, a tapered fuel manifold, a Mark 3H turbopump,
low fuel 6 p injector, and furnace-brazed, stainless steel thrust
chamber. All inboard engines of the Saturn vehicles SA-1 throughSA-202 exhausted turbine gases through exhaust fairings on the
tail unit thrust structure. SA-203 and subsequent vehicles utilize
inboard engines having integral exhaust ducts that exhaust turbine
gases into the thrust chamber exit flow of the inboard engines.
Throughout the Saturn program, the outboard engines have utilized
aspirators for dispersing the turbine exhaust gases. H-1 engines,rated at 165,000 lbf each, propelled Block I vehicles SA-I through
SA-4 to test the concept of engine clustering and multi-tank propel-
lant container. Engines rated at 188,000 lbf each propelled theBlock II vehicles, SA-5 through SA-10, providing primary boost
phase for orbiting boilerplate Apollo payloads and testing the S-IV
stage. Engines rated at 200,000 lbf propelled SA-201 through SA-
205. These engines provided the primary boost phase for testing
the S-IVB stage, IU, and Apollo payloads which were on the SaturnV lunar vehicles. On the Skylab vehicles, the H-I engines are rated
at 205,000 lbf.
H-I Engine Predicted Performance.
The H-1C and H-1D engines meet the requirements of H-1 Engine
Model Specification R-1141dS. The engines must produce 205,000
-*-2,000 lbf, an instantaneous impulse of 261 sec (min) and 263.4
sec (nom), a nominal chamber pressure of 652 psia, and a mixture
ratio of 2.23-----2 percent (O/F). The nominal fuel inlet pressure is
57 psia prior to ignition. Required NPSH during flight is 35 ft.
The minimum required lox pump inlet pressure prior to ignition
Figure 4-9 (Sheet 2 of 2) 4-9
SectionIV S-IB Stage
is 80 psia with a required NPSH for flight of 35 ft. The enginemust have an effective duration of 155 sec (min). Figure 4-10
presents the S-IB-6 propulsion predictions using static test data
and Rocketdyne acceptance test data. These predictions may changeslightly as parameters for individual missions are defined. Predicted
inboard engine cutoff time is 2 min 17.7 sec into flight with the
outboard engines cutoff time occurring 3 sec later. The reliability
assessment of the H-l engine during the AS-206 Design Certifi-cation Review was 0.99 at 0.869 confidence. Refer to Section IIfor vehicle performance.
Loading Limitations.
The engine and its structural mounts, while meeting the gimbalingrequirements, must operate without deformation or failure under
the following conditions: (l) flight loading 8.0 G parallel to the
direction of flight and 0.5 G perpendicular to the direction of flight;(2) flight loading 4.0 G parallel to direction of flight and 1.0 G
perpendicular to flight direction. The engine is designed to with-stand a minimum of 1.5 times the forces resulting from all combina-
tions of the above loading conditions or 4.0 G handling loadsapplied in any direction.
H-1 Engine Servicing.
To prepare the engines for firing, the fuel additive blender unit
(FABU) is serviced with extreme-pressure additive ST0140RB0013,
the thrust chamber fuel jacket is serviced with RP-1 fuel, and the
installation of the LPGG squibless igniters, the SPGG, the SPGG
initiators, the Conax valves, and the hypergol cartridges is com-pleted.
FABU Fill. A portable unit services the FABU with 105 in? (min)
of extreme-pressure additive ST0140RB0013 per engine. The ex-treme:pressure additive, maintained at 120_10 ° F, enters the
FABU through a quick-disconnect on the base of each unit. A
plunger-type indicator on top of the FABU extends when the unitis full.
Thrust Chamber Fuel Jacket Fill. At T-l day, 15 hr, each thrust
chamber fuel jacket is prefilled with 12 ±0.5 gal of RP- l fuel. Filling
each chamber reduces the time delay of fuel entering the combus-tion chamber after the main fuel valve opens. Ambient RP-I enters
the thrust chamber through a quick-disconnect on the GG fuel
bootstrap line.
H-I Engine Ordnance Devices Installation. Installation of ordnance
devices begins prior to RP-I tanking, with the Conax valves in-
stallation, followed on T-l day by installation of the squiblessigniters, solid propellant gas generators, and SPGG initiators. Prior
to installation the components receive checkouts for insulation
resistance, pin-to-case checks, visual inspections, and electrical
checks of the bridge-wires in the initiators and Conax valves. The
harnesses for the elect(o-explosive devices also receive electrical
Checkout for stray voltages and for continuity. The SPGG initiator
circuits are tested while the ignition switch on the S-IB networks
panel is first, in the ARM position, and second, in the SAFE
position. Sixteen squibless igniters, two per engine, are installed
in the LPGG combustorsjust below the injector. Eight solid propel-
lant gas generators, one per engine, attach to the LPGG combustor
SPGG attach flanges. Two initiators for redundancy are installed
in each SPGG. Eight Conax valves, one per engine, attach to the
MLV closing control manifolds located on the main fuel valves.
After checkout of the harness assemblies with no discrepanciesnoted, the electrical connectors are connected to the Conax valve
(two per valve), squibless igniters, and SPGG initiators. Two addi-
tional electrical connectors are connected to the position indicatoron each Conax valve. These connections do not interface with the
ordnance charges.
Hypergol Cartridge Installation. Installation of the hypergol car-
tridge in the engines occurs on T-l day. The cartridge containing
4-10
6 in? of triethylaluminum is inserted into the hypergol container
on the injector. A lockpin secures the cartridge in place. The HY-
PERGOL INSTALLED detector switch senses cartridge installation
and provides a corresponding signal to the S-IB firing prep panel.
O-ring seals on the cartridge prevent leakage of fuel during engineoperation. Burst diaphragms contain the hypergol until ruptured
by fuel pressure during engine start.
Engine Purges.
Four ambient GN 2 purges supplied by valve panel no. 10 in the
ML prevent accumulation of contaminants in engine components.The purges initiated during countdown continue until overcome
by engine internal pressures during start.
Lox System Bypass Purge. Lox system bypass purge, required
anytime thrust chamber exit covers are removed from the engineswhen the stage is on the launch pad, prevents entrance of con-
taminants into the lox dome from the thrust chamber. GNz, froma common manifold that supplies the purge to all engines, enters
the lox dome through the unitized check valve and the heat ex-
changer lox supply line. The purge flowrate is 5(+1.5, -3.3) scfm
per engine at an engine interface pressure of 2 to 8 psig. A poppet-
type check valve in the unitized check valve of each engine prevents
lox flow into the manifold during engine operation. Switch S19,
LOX DOME BYPASS PURGE, on S-IB firing prep panel initiates
the low flowrate purge. Feedback from the lox dome bypass valvein valve panel no. 10 illuminates the LOX DOME BYPASS OPEN
indicator on the S-IB firing prep panel. Pressure switches upstreamfrom the lox dome bypass valve illuminate the LOX DOME
PURGE ON indicator on the S-IB firing prep panel and the firing
panel. Pressure switches on the lox dome bypass supply line illumi-
nate the LOX DOME LOW PRESS. indicator on the S-IB firingprep panel.
Lox System Purge. At T-28 sec, the launch sequencer initiates a
high flowrate lox system purge, 100-+-20 scfm per engine at an
engine interface pressure of 95 -4-10 psig, which is superimposed onthe lox system bypass purge. The purge prevents contaminationof the lox dome from a fuel-rich cutoff in the thrust chamber if
abort is initiated after engine ignition but before liftoff. Lox system
purge can be initiated manually by the LOX DOME PURGE switch
on the S-IB firing prep panel.
Gas Generator Lox Injector and Thrust Chamber Fuel Injctor Purges.
At T-28 sec the launch sequencer also initiates the gas generator
(GG) lox injector manifold purge and the thrust chamber (TC)
fuel injector purge. The GG lox injector purge prevents solid
propellant gas generator combustion products from entering the
lox injector manifold during engine start. The purge prevents con-tamination of the gas generator lox injector and controls turbine
inlet temperature spikes if abort is initiated after engine ignitionbut before liftoff.
Purge flowrate is 76-+-7 scfm per engine at an engine interface
pressure of 255-----25 psig. Lox pressure buildup in the manifold
terminates the purge. The TC fuel injector purge prevents lox from
entering the injector fuel ports during engine start. Flowrate is
238-+35 scfm per engine at an engine interface pressure of
487.5±62.5 psig. Increasing fuel pressure in the injector, as a result
of turbopump acceleration, terminates the purge. The purge pres-
sure must be vented prior to initiating an abort after engine ignitionbut prior to liftoff.
Poppet-type check valves at each engine prevent reverse flow
through the two purge manifolds when engine operating pressuresexceed the purge pressures.
Purge Controls. GAS GENERATOR LOX INJECTOR PURGE
and THRUST CHAMBER FUEL INJECTOR PURGE switches
on the S-IB firing prep panel permit manual actuation of the
purge. Pressure switches on the purge supply lines in valve panel
NOM 205.00 263.63 536.86 240.74 704.71 2,2300 6,691.3
1 206.00 262.72 541.66 242.43 709.32 2.2343 6,754.3
2 205.33 262.82 540.01 241.26 709.27 2.2383 6,690.8
3 205.49 262.82 541.54 240.30 710.03 2.2536 6,687.8
4 205.68 262.39 541.71 242.19 705.52 2.2367 6,749.9
5 206.27 263.26 542.69 240.85 709.08 2.2532 6,771.1
6 205.47 262.85 541.41 240.30 706.45 2.2530 6,694.4
7 205.49 263.55 541.18 238.51 703.84 2.2690 6,696.7
8 205.17 263.28 539.91 239.39 701.11 2.2554 6e674.0
AvG _ I_> I_ I_205.61 262.96 541.26 240.65 -- 2.2491 --
[_:> THRUST ALONG LONGITUDINAL AXIS
J_ INCLUDES FUEL USED AS LUBRICANT
AMBIENT PRESSURE 14.7 PSIA, LOX DENSITY 70.79 LBM/FT 3, LOX PUMP
INLET PRESSURE 65 PSIA, FUEL PUMP INLET PRESSURE o'7 PSIA
CH-_0_
Figure 4-10
no. l0 provide feedback signals to indicators, GO LOX INJECTOR
PURGE and TC FUEL INJECTOR PURGE ON, on the S-IB
firing prep panel and the firing panel. Each of the four purge
switches on the S-IB firing prep panel has three positions: OFF,
AUTO, and ON. The LOX DOME PURGE, GG LOX INJECTOR
PURGE, and TC FUEL INJECTOR PURGE switches must be
in the AUTO position for final countdown as a "prerequisite for
PURGES ARMED signal, an interlock for FIRING COMMAND.PURGED ARMED is monitored on firing panel. The ALL EN-
GINES RUNNING signal closes the TC fuel injector purge valve
and opens the supply regulator dome vent valve to effect regulator
closure. Redundant vent valves in the thrust chamber purge vent
valve panel open, venting the 550-psig purge supply line to atmo-
sphere before umbilicals disconnect. The COMMIT signal closes
the lox system purge valve and the GG lox injector purge valve
by removing + 1D116 bus power. In the event of an abort before
liftoff, the cutoff command reinstates the lox system purge and
GG lox injector purge to prevent vapors from entering the lox
components during the fuel-rich shutdown. The lox system purge
is required for a minimum of 15 sec after engine cutoff, then the
lox system bypass purge continues until the thrust chamber exit
covers are reinstalled. The GG lox injector purge is required for
a minimum of 10 min after engine cutoff. Supply pressures and
flowrates are the same as preflight requirements.
Engine Gearbox Pressurization and Lox Pump Seal Purge. Engine
gearbox pressurization and purge, and lox pump seal purge begin
with pressurization of the S-IB stage control pressure system and
continue throughout propellant tanking, engine start, launch, and
powered flight. GN 2 flowing into the gearbox provides a constant
purge during countdown. During engine operation the GN 2 pres-
surizes the gearbox to prevent lubricant foaming at high altitudes
and to hold the gearcase lube drain relief valve open precluding
flooding of the gearcase with lubricant. GN 2 from the 750-psig
control pressure system is orificed to a 1.85(+ 0.6, -0.8) scfm per
engine flowrate, which maintains the gearbox pressure between
2 and 10 psig. The lox pump seal purge, applied in the area between
lox and lube seals, keeps the lox and lube seals leakage separated
to prevent formation of a highly explosive hydrocarbon gel in the
turbopump. The GN 2 pressure forces any seal leakage overboard
Section IV S-IB Stage
through separate drain lines. The lox pump seal purge supply tees
off the gearbox pressurization line and is orificed to purge thelox seal at a 1.35+0.6 scfm per engine flowrate. In the event of
an abort before liftoff the lox pump seal purge is required to
maintain separation of lox and lubricant leakage and continues
until the pump returns to the ambient temperature. The gearbox
pressurization purge supply maintains gearbox pressure between
2 and 10 psig.
H-1 Engine Operation.
At T-3 sec, the ignition sequencer issues the IGNITION com-
mands to start the engines in pairs, opposite inboard engines
5 and 7, and 6 and 8; and opposite outboard engines 2 and 4,
and 1 and 3. The engines start in this order, 100 msec apart per
pair, to reduce stress on the vehicle during engine start. The elec-trical command fires the redundant SPGG initiators on each SPGG.
See figures 4-11 and 4-12. The SPGG propellant begins burning,
producing hot gases that accelerate the turbine and ignite the two
liquid propellant gas generator squibless igniters. Through a gear
train, the turbine drives the turbopump causing discharge pressures
to increase against the closed propellant valves. During tanking
operations fuel and lox filled the suction lines, the pump sections
n IGNITION SIGNAL
TURBOPUMP
FUEL CONTROL PRESS. INCREASES TO 300 PSIG
MAIN LOX VALVE OPENS
IIGNITER FUEL VALVE OPENS
HYPERGOL FLOWS
IGNITION FUEL FLOWS
LEGEND
n OPERATION STARTS
R OPERATION ENDS
_ OPERATION CONTINUES
VA LVE ACTUATIO N TIMEINDICATED POSITION MAINTAINEDI I I
I i i l
0 0.25 0.5 0.75
IGNITION
JJJ_ IGNInON MONITOR VALVE OPENS
JJ_Jj LPGG CONTROL
OPE7LPGG OPERATES
(BOOTSTRAP)
] " N_Ss TEADYTA_E
1 1.25 SECO ND S
C-H 20057
Figure 4-11
4-11
Section IV S-IB Stage
H-I ENGINE START AND RUN OPERATION
FUEL
LOX
LUBRICANT
GAS GENERATOR
EXHAUST
HYPERGOL
THRUST CHAMBER
PRESSURE
ELECTRICAL
MECHANICAL
LINKAGE
LEGEND
Ilnlilnlnnlmmliln
mmmmunmiummmn
illUlllllnlUlllllllllllilllllUlUl|
mi mmm
roll i OK
FABU-
GEARBOX
PRESSURIZATION -
X SEAL IIPURGE --
(_) IGNITION COMMAND AT T-3 SEC FIRES SOLIDPROPELLANT GAS GENERATOR (SPGG) INITIATORS. i
(_ SPGG IGNITES PRODUCING GASES TO DRIVE iTURBOPUMP TURBINE AND IGNITE SQULBLESS I
IGNITERS. i(_) TURBINE STARTS FUEL AND LOX PUMPS,
(_ PUMPS DELIVER PROPELLANTS UNDER PRESS.TO MAIN VALVES. i i. • ,S
r ....... + ...... n-'" ii ' i
(_ (FASU) AND MIXES WITH EXTREME-PRESSURE I iFUEL ENTERS FUEL ADDITIVE BLENDER UNIT . _ _ __
ADDITIVE. MIXTURE LUBRICATES TURBOPUMP I • _ --_
GEARTRAN - • - •(_E) CONTROL PRESS. FUEL OPENS MAIN LOX t I hb'_L_/"_'_7_ _ --_
VALVE (MLV). LOX FLOWS INTO ENGINE I m_<_Aq) - -_.
COMBUST*ONCH BER. ' ,ll1 Jk --= --(_ MLV MECHANiCALLY OPENS IGNITER FUEL l MLV'_ • _ ._'_._ \ E m
VALVE (IFV) WHEN MLV OPENS 50-70 DEG. " J_ A \ • viv_ " LMEV ==CONTROL PRESS. FUEL FLOWS TO HYPERGOL " "% _'o1,-,'_-_\ • -- _ • ECONTAINER AND IGNITION MONITOR VALVE n . _j\ • • : • . . . ;,NLm i " _/._-_L_ i I - i T i -
(_ FUEL BURSTS TWO DIAPHRAGMS IN HYPERGOL _ : %'-- _2_-_--_-,_'_ -- / -- _ " " Z E
CARTRIDGE. HYPERGOL FLOWS THROUGH _"_1 _l=#_m, _, _ _ll' -- '_1 -- ' I1_ ' _ I -- I i ..... i .....FUEL INJECTOR INTO COMBUSTION CHAMBER. V_L,-. _,_ • IMV-_,. _
0 IGNITION. HYPERGOLANDLOXIGNITE _ '_r_; ml_ "_J i th 1i" |
"-' _' _i:"n ..... m' _ -=® TcFUEL_NJECTORMAN,FOLDPRESSURERESULT,NG_ _." -- "_ _%'Y'/I_ = = 'FROM INCREASED COMBUSTION CHAMBER PRESS. / _1 _ _ " | • I _ o/• L|JW _ ESHUTTLES IGNITION MONITOR VALVE (IMV) AL- / L _ | l • _i_ _ _ _ • i_
LOWING FUEL TO OPEN MAIN FUEL VALVE (MFV}. / i I _l_ I I ,_1 .... • .... _Jl_' _ _ =''% E
,-,FV i I m • s i -= II_l//__- -_1 I _•(_ FUELFLOWSTHROUGHMEVINTOTHRUSTCHAMBER | • i • • _.,-:_-L : •FUELJACKETANDTHROUGHFUELINJECTORINTO J- u • - / i -_: o_1 _ ; •COMBUSTION CHAMBER TO SUSTAIN IGNITION _ U • • / " •_ _fJ, i I
M ' -_ _- n • i '41 =-(_ TC FUEL INJECTOR ANIFOLD PRESSURE RESULT- _i__ _ • -- 1 _ _ " " t • I_ •
ING FROM INCREASED COMBUSTION CHAMBER k'_ "_ _1_ I • I I i_i/. • ._ I _ _ --'_
PRESS OPENS LIQUID PROPELLANT GAS GENERATOR _ _-I _'i \ | = _--" /_l/ • _-- I I I •
(LPGG) CONTROL VALVE. BOOTSTRAPFUELAND I @ \ I _ _ /_i 5"" _. i •LOX ENTER LPGG COMBUSTOR. SPGG GASES AND n .v;,o_,_\ , _ J • - _E LPGG _ I ; •
SQUIBLESS IGNITERS IGNITE BOOTSTRAP PROPELLANT n r_'_T_,_F_ | n " =1 _ = CONTROL VALVE ..] I I •• . CONTAINER I • I _- _ = I _ "_
0 SQU*BLESS,ON.E._URNAFPROX_SECTOI n • _ H /' - i . --=ENSUREBOOTSTRAPPROPELLANTIGNITION. I I _ ¢_ _ e I e -- t I. --
O LPGGGASESCAUSETURBOPUMPTOACCEL - l ,iI'l _ - ;: -ERATE TO OPERATIONAL SPEEDS. 1,_,_j=. m _,_I_------_ _ _ I _I_ GOX_p--_ = I |
(_ THRUST OK PRESS. SWITCHES (TOPS) SENSE FUEL II.'__: t ._--.._ _ ---
PRESS. AND SEND THRUST OK SIGNALS TOIU _.-_i___ _ S_i_ _ _-'_ _
EDS DISTR VOTING LOGIC CKTS THAT GENERATE /// ¶-'_._/-_ _-_ _'_ • K_I er"_5"_
'ALL ENGINES RUNNING' SIGNAL & A COMMIT J:_ I l_aj_Im ,Ha _ _, \-_1 / _I_F#F'-;_ _._'__--r J
,NTERLOCK.LOG,CCKFOUTPUTSOFERATE _ / _9%/ _ _/ __.C_' SQU,_LESS--//\ _ I'L/V ENGINES' LIGHTS ON COMMAND MODULE _ _ / _ _ IGNITERS.-/_ dlI_"_ _
EIGHT 'L/V ENGINE' LIGHTS EXTINGUISH AS RE- THRUST _ _\ _m LPG_ j
® SPECT,VEENG,NESATTA,NSAT,SFACTORYTHRUBT.CHAMBER-- (_ _ " HEATHEAT EXCHANGER CONVERTS kOX TO GOX _ I _ EX_ A.,_.
FOR S-,B LOX TANKS INFUGHT PRESSURIZATION. _ "-_":..':iii _ ..... R SPGG--_.._
ASPIRATORS.
(_ LOX SEAL PURGE AND GEARBOX PRESSURIZA-
TION FROM CONTROL PRESS. SYS CONTINUES
THROUGHOUT S-IB STAGE OPERATION. o g / INITIATORS CH-T4212-2
Figure 4-124-12
of the turbopump, and both high-pressure propellant ducts down
to the closed gates of the main fuel and lox valves. Control-pressure
fuel tapped off the high-pressure fuel duct flows through a series
control line to the main lox valve (MLV), igniter fuel valve (IFV),fuel additive blender unit (FABU), and to the Conax valve. When
fuel pressure increases to 70-150 psig the FABU opens to permit
lubricant flow to the turbopump bearings and gears. As the tur-
bopump accelerates and fuel pressure increases to 300±50 psig,
control-pressure fuel opens the MLV permitting lox to enter the
lox dome and gas generator lox bootstrap line. A cam on the MLV
gate shaft mechanically opens the IFV when the MLV has traveled
50 to 70 degrees open, which permits fuel flow to the ignition
monitor valve (IMV) and to the hypergol cartridge. Inlet and outlet
burst diaphragms in the hypergol cartridge break at 300+---25 psig
and the control-pressure fuel forces the pyrophoric fluid through
seven igniter fuel ports in the injector into the combustion chamber.The hypergol ignites spontaneously with lox entering the combus-
tion chamber through the injector lox ports. Propellant ignition
causes a pressure increase in the combustion chamber, which is
sensed by the IMV. The IMV shall not open at 12+-0.3 psig, and
must open at 22-+-0.3 psig. Nominal 1MV opening pressure is 15
psig. The 1MV permits control-pressure fuel from the IFV to flowto the main fuel valve (MFV) opening actuator resulting in MFV
opening. The pressure required to open the MFV is 350±50 psig.
Fuel, under turbopump pressure, flows through the MFV to the
gas generator fuel bootstrap line, through the thrust chamber fuel
jacket, through the injector manifold, and into the combustion
chamber where main propellant ignition occurs. Increasing fuel
pressure resulting from main propellant ignition and sensed at the
thrust chamber (TC) fuel injector manifold opens the liquid propel-
lant gas generator (LPGG) control valve. The control valve fuel
poppet opens at 105+---20 psig; the lox poppet opens at 200±20
psig.
Bootstrap propellants enter the LPGG with a slight lox lead and
are ignited by SPGG hot gases and the two squibless igniters. Gasesfrom the LPGG cause the turbopump acceleration to continue until
rated thrust is attained. Calibrated orifices in the fuel high-pressure
duct inlet and in the LPGG lox and fuel bootstrap lines control
the propellant flow, thus allowing the engine to operate at ratedthrust.
Three thrust OK pressure switches (TOPS) on each engine sense
fuel pressure downstream from the MFV to indicate satisfactory
engine thrust. The switches actuate at 785-+-15 psig, indicating satis-
factory thrust attained. TOPS outputs are inputs to voting logiccircuits in the IU EDS distributor. The EDS distributor provides
the ALL ENGINES RUNNING signal, when all engines have
attained satisfactory thrust. Engine running indications monitored
on LCC panels, are also monitored on the main display console
(MDC) in the command modules. The MDC L/V ENGINES lights
extinguish when the engines attain rated thrust. Gearbox pres-surization and lox pump seal purge, supplied by the S-IB stage
control pressure system continues throughout flight. Sensors in-
stalled on each engine provide information on engine conditions
during flight to the S-IB stage telemetry system. See H-1 Engine
Measuring. See figures 4-13 and 4-14 for H-1 engine lines andorifice summaries.
H-1 Engine Cutoff.
The launch vehicle digital computer (LVDC) issues the engine
cutoffcommands in two steps through the S-IB stage switch selector.
The first command, issued approximately 2 min 17.7 sec into flight,shuts down the four inboard engines. Approximately 3 sec later,
the second command shuts down the four outboard engines. A
Conax valve on each engine effects engine cutoff. See figures 4-15and 4-16. The cutoff command fires redundant explosive actuators
which open the Conax valve by sheering redundant metal dia-
phragms in the valve body. Control-pressure fuel flows through
Section IV S-IB Stage
H-i ENGINE LINES SUMMARY
LINE
FUEL H. P. DUCT
LOX H, P. DUCTMFV CONTROL
MLV OPENING
MLV CLOSING
IGNITER FUEL
IMV SENSING
LUBE FUEL INLET
FABU DISCHARGE
GG & IMV CONTROL
FUEL BOOTSTRAP
LOX BOOTSTRAP
TOPS SENSING
TURBINE EXHAUST
HEAT EXCHANGER INLET
(LOX)THRUST CHAMBER
FUEL
LOX
GEARBOX PRESSURIZATION
AND LOX PUMP SEAL
PURGE
SIZE
(IN. DIA.)
3.25 ID
3.375 ID
1/4i/2
3/8 0D)3/8i/41/23/s3/8
11/4
3/41/4
3/4
1/4
FLOWRATE
(NOMINAL)
241.0 Ib/sec
537.4 Ib/sec
1.4 Ib/sec-.
0.63 Ib/sec
13.52 Ib/sec
4.61 Ib/sec
18.13 Ib/sec
3.0 Ib/sec
227.5 [b/sec
532.8 Ib/se c
3.2 =cfm
Figure 4-13
CH.14271.1
H-1 ENGINE ORIFICE SUMMARY
ORIFICE
FUEL DISCHARGELOX BOOTSTRAP
FUEL BOOTSTRAP
GEARBOX PRESSURIZATION
LOX PUMP SEAL PURGE
HEAT EXCHANGER (3)
FUEL BLEED
MLV OPENING
FABU OUTLET
LOX BOOTSTRAP (FIXED)
SIZE
(IN. DIA)
2.680
0.356
0.700
0.013
FL OWRATE OF
1.7 scfm AT
750 pslg, 70°F
0.101
0.060
0.116
0.147
0.4O0
J_ ENGINE CALIBRATING ORIFICES
(NOMINAL SIZE)
CH.14269.1
Figure 4-14
the Conax valve to the MLV closing actuator. MLV closure stops
lox flow to the combustion chamber and the LPGG causing thrust
chamber pressure and turbopump speed decay. Closing the MLV
first in the cutoff sequence permits a fuel-rich cutoff to prevent
a temperature spike in the LPGG and thrust chamber. The MLV
permits the IFV to close mechanically and shut off control-pressurefuel to the IMV and hypergol inlet. The MFV closes by spring
tension when its actuation pressure decays below 145(+40, --50)
psig, shutting off fuel flow to the thrust chamber and LPGG. Fuel
pressure decay causes TOPS deactuation, which closes the pre-valves and removes the thrust OK indications to the IU EDS voting
logic circuits. Loss of thrust OK" signals causes the L/V engines
lights on the MDC to illuminate, indicating to the astronauts that
the engines have cut off. L/V ENGINES -5, -6, -7, and -8
lights will illuminate first, indicating inboard engines cutoff. Ap-
proximately 3 sec later, L/V ENGINES --1, -2, --3, and -4
lights will illuminate, indicating outboard engines cutoff. The lights
will remain on until S-IB/S-IVB separation. When fuel pressure
in the fue! injector manifold decays below 105"+-20 psig the LPGG
control valve closes by spring pressure. The pressure decay also
permits the FABU to close, shutting off lubricant flow to the
turbopump gearbox. The FABU valve closes when the inlet pressure
is 50-90 psig. Engine thrust decays to zero in approximately 3.5
to 4.5 sec after receipt of cutoff command. During flight, if an
4-13
SectionIVS-IBStageengine malfunctions and turbopump discharge fuel pressure falls
between 790 and 740 psia, TOPS deactuation on that engine will
initiate engine cutoff. See Electrical Sequencing for additional
information on engine shutdown.
H-1 Engine Cutoff Commands.
During flight, any of three basic cutoff modes will initiate H-I
engine cutoff. Those modes are: malfunction cutoff, normal 4-by-4
cutoff, and range safety command cutoff. The normal 4-by-4 cutoff
is the planned engine cutoff mode, where the inboard engines shutdown, followed by outboard engine shutdown 4 sec later. The
malfunction mode will shut down individual engines while the range
safety command will shut down all engines simultaneously to effecta zero-thrust condition. The range safety commands originate from
the ground stations and are used in the event of vehicle deviation
from planned trajectory or other emergencies endangering the
launch facility or the mission. All other cutoff commands originatein the vehicle.
Normal 4-by-4 Engine Cutoff. The switch selector starts the engine
cutoff sequence by enabling the four propellant level sensors at
2 min I 1.2 sec into flight (times are predicted S-IB-6 flight times
which may change as mission parameters are defined). See figure4-18. One of the sensors will actuate at 2 min 14.7 sec (T2) and
provide an input to the LVDA/LVDC to start the cutoff sequence.At 2 min 17.5 sec (T 2 + 3.0 sec) the LVDC will issue the inboard
engines cutoff command through the S-IB stage switch selector.
The cutoff command fires both squibs in the Conax valves on
the inboard engines to start the cutoff operation. See H-I Engine
Cutoff. At 2 min 19.0 sec (T 2 + 4.5 sec), a switch selector command
groups together the outputs from the voting logic of the thrust
OK pressure switches on each outboard engine. If any outboard
engine experiences lox starvation and shuts down, deactuation of
the TOPS on that engine will command the remaining three engines
H.i ENGINE CUTOFF SEQUENCE
LEGEND
OPERATION _J_ OPERATIONSTARTS CONTINUES
lj_ VALVE ACTUATIONOPERATION TIME INDICATEDENDS POSITION MAINTAINED
THRUST DECAYS TO ZERO IN 3.5 TO 4.5 SECONDS
I I IJ ENGINE CUTOFF SIGNAL
I I_J_ CONAX VALVE RRES
7'. I I LI FUEL PRESSURE ON MAIN
LOX VALVE CLOSING PORT
,MA N LOX ViLVE CLOiES
J_ FUEL IGNITERVALVE CLOSES
J_j_ _[_ GG CONT ROL VALVE CLOSES
m I MA,NFUE II] [.,"VALVECLOSES
iI I rHRUSTDECAYS,Y,0_ [_• THRUST
DECAYS
• BY 100%i
0 0.25 0.5 0.75 1.0 1.25 1.5 1.75 2.0
SECONDS
C½.20058
4-14
Figure 4_15
to cut off. At 2 min 19.5 sec (T 2 + 5.0 sec) the switch selector
enables the fuel depletion sensors. This command permits the fuel
depletion sensors to issue the cutoff command if the fuel drops
below the sensor level in the tank sumps to prevent the engines
from shutting down because of fuel starvation. The engines must
shut down with a fuel-rich mixture to prevent excessive temper-
atures in the thrust chamber and gas generator. Either lox starvation
or fuel depletion cutoff will establish the vehicle staging time base
(T3). The switch selector issues a backup cutoff command at 2
min 21.7 sec (T s + 0.1) to ensure engine shutdown for staging.
Simultaneous cutoff of all outboard engines is necessary to prevent
attitude deviations that would endanger the launch vehicle during
staging.
One Engine-Out Capability. The S-IB stage has no engine-out
capability during the first 3 sec of flight to preclude the possibility
of a catastrophic vehicle/tower collision. At 3 sec into flight (T_
+ 3 see), the switch selector initiates a command to enable a
one-engine-out capability. Failure of one engine, resulting in TOPSdeactuation, will initiate cutoff for that engine. Shutdown of that
one engine will immediately remove bus power from the TOPS
voting logic circuits of the other engines, thereby inhibiting the
cutoff circuits for the remaining seven engines regardless of their
performance (figure 4-17). This one-engine-out capability will re-main in effect until 10 sec into the flight (T 1 + 10 sec) to ensure
maximum available thrust to lift the vehicle clear of the pad area.
Multiple-Engine-Out Capability. Ten sec into flight (T I + 10 sec)
the switch selector disables the one-engine-out bus (figure 4-17),
which in turn enables the bus power to the TOPS voting logic
circuitry. If an engine (or engines) fails, thrust decay in that enginewill cause TOPS deactuation, which fires the Conax valve to com-
plete cutoff for that engine. For the remainder of flight, until TOPS
grouping command, each engine can shut down independently.
Range Safety Cutoff. Any time during flight that the vehicle deviates
beyond acceptable limits of the intended trajectory or becomes
a hazard, the range safety officer (RSO) can destroy the vehicle.
As a requirement of the Air Force Eastern Test Range, thrust of
liquid propellant vehicles must be reduced to zero before destroyingthe vehicle. The R.SO issues two commands to destroy the vehicle.
The first command shuts off all engines simultaneously and also
charges the propellant dispersion system EBW firing units. The
second command triggers the firing units to detonate the explosive
shaped charges (See Ordnance). The first command fires the squibs
in all Conax valves simultaneously, thereby initiating eight engine
cutoff (figure 4-17). Thrust decays to zero within 3.5 to 4.5 sec.
Thrust Chamber.
The thrust chamber consists of a gimbal assembly, an oxidizer
dome, an injector and hypergol container, and a thrust chamber
body. The oxidizer dome, installed over the injector, encloses theforward end of the thrust chamber. The thrust chamber receives
propellants under turbopump pressure, mixes and burns the pro-
pellants, and imparts a high velocity to the expelled combustion
gases to produce thrust.
Gimbal Assembly. The gimbal assembly, mounted on the oxidizerdome, secures the thrust chamber to the thrust pad of the tail
unit assembly. The gimbal is essentially a universal joint mounted
on thrust vector alignment slides. The gimbal assemblies on the
inboard engines are functional only to the extent of aligning the
thrust vector during engine installation. The outboard engine gim-
bal assemblies permit the hydraulic actuator to change the thrust
vector to satisfy the guidance and attitude corrections. The S-IB
stage outboard engine gimbal assemblies permit an 11.31-deg an-gular displacement of the geometric thrust vector from the normal
plane of the gimbal bearing axis (with both actuators fully extendedor retracted, either in-phase or out-of-phase). A single actuator,
Section IV S-IB Stage
1-1ENGINECUTOFFOPERATIOI
FUEL
LOX
LUBRICANT
THRUST CHAMBER
PRESSURE
ELECTRICAL
MECHANICAL
LINKAGE
(_) CUTOFF SIGNAL FIRES EXPLOSIVE ACTUATORS INCONAX VALVE. CONAX VALVE OPENS AND
PORTS CONTROL-PRESS. FUEL TO MLV CLOSINGACTUATOR.
(_MLV CLOSES. FUEL PRESS. OVER LARGER AREAPLUS SPRING OVERCOMES OPENING PRESS. LOX
FLOW TO THRUST CHAMBER AND GAS GENERATOR
TERMINATES. MLV CLOSURE PERMITS MECHANICAL
CLOSURE OF IFV.
(_IFV SHUTS OFF FUEL PRESS TO IMV AND HYPER-GOL CONTAINER.
(_) THRUST CHAMBER PRESSURE DECAYS (NO LOXTO SUPPORT COMBUSTION OF FUEL).
(_) IMV CLOSES AS THRUST CHAMBER PRESS.DECAYS.
(_ LPGG VALVE CLOSES AS THRUSTCONTROL
CHAMBER PRESS. DECAYS. LOX POPPET
CLOSES FIRST.
(_ FUEL-RICH SHUTDOWN OF LPGG PREVENTSEXCESSIVE TEMPERATURES IN COMBUSTOR.
(_ TURBINE AND TURBOPUMP SPEED DECAYS AT
LPGG SHUTDOWN. FUEL AND LOX PUMP DIS-
CHARGE PRESSURES DECAY.
(_MEV OPENING PRESS, BLEEDS OFFTHKE)UGHIMV OVERBOARD DRAIN LINE AND VALVE
CLOSES. MFV REMAINS OPEN UNTIL THRUST
CHAMBER PRESS. DECAYS TO EFFECT A FUEL-
RICH SHUTDOWN.
(_ FUEL PUMP DISCHARGE PRESS. DECAY CAUSESGEARBOX LUBRICANT FLOW TO TERMINATE.
(_ FUEL PUMP DISCHARGE PRESS. DECAY CAUSES
TOPS DEACTUATION.
(_ L/_/ ENGINES -5, -6, -7, & -8 LIGHTS ILLUMI-NATE INDICATING INBOARD ENGINES CUT-
OFF. L/V ENGINES -I, -2, -3, & -4 LIGHTS
ILLUMINATE INDICATING OUTBOARD ENGINES
CUTOFF.
(_ TOPS DEACTUATION CLOSES PREVALVES.
THRUST
SQUIBLESS
Figure 4-16 4-15
SectionIVS-IBStage
H-I ENGINE CUTOFF LOGIC
J_ ENGINE-OUT LOGIC
COMMAND MULTIPLE ENGINE _ +1D16CUTOFF ENABLE T+10 SEC. _J_
COMMAN0ENG. _JT_--J=__,I _ JNO._CUTOPP(+;Dn) I
t _t L-'_J_"/ J__ POWER I-.IN _
TRANSFER _O_'
INHIBIT a J
÷1034 BUS
POWER
COMMAND ENG. _ ._
NL. 8 CUTOFF (+1Dll) _
ENGINE CUTOFF LOGIC (TYPICAL ALL ENGINES)
SWITCH NO. 1
THRUST O.K.
SWITCH NO. 2
SWITCH NO. 2
THRUST O.K.
+1Dll =1_ NO.3=_
LIFTOFF _JO_'
NOTES COMMAND SINGLE ENGINE CUTOFF ENABLE TI +3 SEC _L 7
, T,MEsSHOWNARE ,r_-_L_--T..... --_W,TCHTYPICAL FLIGHT TIMES J I J A- +1D11 THRUST O.K, J
2 COMMANDS ARE ISSUED BY i j I fTHE S-IB STAGE SWITCH SELECTOR CLOSE 141_1_ SWITCH NO. 1
'_NOENG,NEOOTLOGC_R_l COMMANDF_ THRUSTOKFIRST 3 SEC OF FLIGHT: ONE J J
ENGINE-OUT LOGIC FROM 3 SEC I -- " COMMAND ENGINES
COMMAND ENG.
NO. I CUTOFF
(+1Dli +1D21)
POWER
TRANSFER (+1D11)
TO 10 SEC; MULTIPLE ENGINE
OUT CAPABILITY ENABLED 8Y
SWITCH SELECTOR AFTER 10 SEC
OF FLIGHT.
._ENG. NO. I PREMATURE CUTOFF INHIBIT
_ENG. NO. 2 PREMATURE CUTOFF INHIBIT
4_ENG. NO. 3 PREMATURE CUTOFF iNHIBIT
5_ENG. NO. 4 PREMATURE CUTOFF INHIBIT
b_>.TYP I OF 4 PROPELLAI_r LEVEL SENSORPREMATURE SIGNAL INHIBIT CIRCUIT
_TYP I OF2 FUEL DEPLETION SENSOR
PREMATURE SIGNAL INHIBIT CIRCUIT
r- _'_AZ--I s_u,BsNO_SAEEII VALVE' ._JCOMMANDENGJ _ I /" r"F-NO. 1CUTOFF(+1D21)I
J NO. 2 +1Dll
II I __o,_G
J S_UIB J _"_J CO MMAN D ENGINES
' NO.; I Isou,BsNO.,SAFEIL__-
+IDll
THRUST
NORMAL 4 BY 4 ENGINE CUTOFF LOGIC |2 MIN. 17.5 SEC (1"2 +3.0) INBOARD
ENGI NES CUTOFF COMMAND
2 MIN. 11.2 SEC. PROPELLANT
LEVEL SENSOR ENABLE COMMAND
FUELDEPLETIONSENSORSI r,,, _ _ i3"> I I
ENABLECOMMANDFUEL DEPLETION _ 1.1 J'_'_J _r_,m1_-,_
SENSO B ACTUATED IJ I W
TOPS DEACTUATION t_ I J _.TART
(_) RESULTING FROM LOX DEPLETION ,_!_,_ _ _3_3
2MIN. 19.0 SEC (T2 + 4.5 )
TOPS GROUPING COMMAND J J
(LOX DEPLETION ENABLE CMD) J J
PREPROGRAMMED 4 BY 4
(_ /tl ENG. CUTOFF..(._. jL_ COMMAND
_J_ RANGE SAFETY/EDS
CUTOFF COMMAND
FROM 2 OF 3
VOTING LOGIC
ENGINE
i_) CUTOFFLOCKUP
MEASUREMENT_" ENG OUT TO I.U.
4-16
TOPS GROUPING LOGIC
• INBOARD (_ ! (OUTBOARD ENGINES ONLY)
r_ _ ENGINES CUTOFF "_--_ l
I \ I _ (+IDll&+tD2I) "J" l +,0,2
OUTBOARDtl- _ I _ ENGINES CUTOFF I
J 2 MIN 24-7 SEC (T3 +0.1) (+lOll& + 1D',
I OUTBOARD ENGS _
CUTOFF COMMAND ipl O).I_I]I_"
TsTOAI_ _ I r-_
T2 +IDll
Figure 4-17
CUTOFF (+ID11) _ _
CH-20059
®
_.©
either fully extended or fully retracted, permits an 8-deg angular
displacement of the geometric thrust vector. The gimbal assembly
design will permit a 14.8-deg (max) geometric thrust vector angular
displacement using two actuators, or a 10.5-deg (max) displacement
using one actuator. The actuators determine the limits of the angular
displacement for the S-IB stage outboard engines. The alignmentslides on both inboard and outboard engines provide lateral posi-
tioning of the engine geometric thrust vector with respect to the
engine centerline.
Oxidizer Dome. The oxidizer dome directs lox from the turbopump
into the injector, provides a mount for gimbal and thrust chamber
assembly, and transmits engine thrust to the vehicle structure.
Bootstrap lox for the gas generator is supplied from the oxidizerdome.
Injector. The injector (figure 4-18) receives iox and fuel and injects
the propellants into the combustion area in a fuel-on-fuel andlox-on-lox pattern to insure satisfactory combustion. Twenty-one
concentric passage rings distribute the fuel and lox through angled
orifices to attain the like-on-like impingement pattern. Fuel flows
through the outermost ring and in each alternate ring. For increasedcombustion stability, copper baffles mounted on the injector divide
Section IV S-IB Stage
the injector face into six equal areas around a center hub. Passages
in the baffles that correspond to orificed holes in the lox and fuel
rings permit propellant flow through the baffles into the combustionarea. Hypergol flows through seven passages in the injector to fuel
housings brazed into the injector face (one in each baffle compart-ment). Fuel from the thrust chamber body enters the injector
through a ceramic-coated screen located around the injector pe-riphery and lox enters the injector from the forward side, which
is enclosed by the oxidizer dome. The injector transmits thrust
forces, produced by combustion zone pressure acting upon the
injector, to the oxidizer dome and subsequently to the vehiclestructure. See figure 4-19 for injector characteristics.
Thrust Chamber Body. The thrust chamber body is a de Laval
structure (100 percent bell) and consists of a converging combustion
chamber, throat section, and diverging section through which com-
bustion gases are expanded and accelerated. Longitudinal stainless
steel tubes joined together by furnace brazing and retained by
external rings and tension bands form the thrust chamber body.
This type construction permits regenerative cooling during engine
operation by fuel flow through the tubes. A tapered fuel manifoldattaches to the top of the thrust chamber and provides equal flow
H-I iNGINE INJECTOR
SECTION A-A
rl
FUEL
ENTRANCE JSCREEN
BAFFLE ,_
PASSAGES J
i
).... ill /
; ii I I' II II
II IIII II
_HUB
PASSAGES
_ LOXG (10)
rtlllllli llJ7
Figure 4-18
NG
INJECTION
ORIFICES SECTION B-S
IGNITER
BAFFLE
FINS
HUB
C-H 14238
4-17
Section IV S-IB Stage
THRUST HAMBER BODY CHARACTERISTICS• THRUST CHAMBER INJECTOR CHARACTERISTICS
ITEM
MATERIAL
FABRICATION METHOD
PLATE AREA
PLATE DIAMETER
SCREEN MATERIAL
OXI DIZER PRESS. DROP
FUEL PRESS.DROP
ITEM
TOTAL NO. ORIFICES
TOTAL ORIFICE AREA
IMPINGEMENT DISTANCE
iMPINGEMENT ANGLE
NO. FILM COOLANT
ORIFICES
NO. BAFFLE FIN COOLANTORIFICES
NO. HUB COOLANT
ORIFICES
PERCENT FILM COOLANT
DESCRIPTION
347 STAINLESS STEEL CRES WITH
COPPER RINGS
FURNACE BRAZED332 in. 2
20.55 _n.
18-8 cres CERAMIC-COATED
10-mesh (0.059 in. dla)
122.2 psl AT 535.0 Ib/se¢
85.4 psi AT 235.0 Ib/se¢
LOX FUEL
1137 1394
10,182 in.2 7,044 _n.2
0.571 in. 0.300 in.
40 deg 40 deg
0 84
42 42
0 44
0 2.1
CH-14217-2
Figure 4-19
through each down tube. This even flow through the cooling tubesincreases the thrust chamber life. The return manifold encloses
the tubes at the nozzle base and directs fuel flow through up-tubes
to the fuel injector. See figures 4-20 and 4-21 for thrust chambercharacteristics.
Propellant Feed System.
The propellant feed system consists of the turbopump, fuel and
lox high-pressure ducts, main lox valve, main fuel valve, igniter
fuel valve, check valves, and orifices. This system supplies the
propellant, at the prescribed flowrates and pressures, to the thrust
chamber and gas generator.
Turbogump. A turbine-driven, dual-pumping unit turbopump de-
livers fuel and lox at high-pressure and high flowrates to the enginecombustion chamber. The turbopump (figure 4-22) consists of an
oxidizer pump, fuel pump, reduction gearbox, accessory drive
adapter, and gas turbine. The turbopump mounts on the side of
the thrust chamber and requires only two short, high-pressure ductsconnecting the volute outlets to the main propellant valves. This
installation assures minimum pressure drop from pumps to valves.Bootstrap propellants from turbopump discharge flow to the LPGG
for turbine operation. The high-speed gas turbine drives the tur-bopump through a series of reduction gears that drive the main
shaft. In each pump, axial-flow inducers increase the pressure at
the impeller inlet (thus requiring a low NPSH), and radial-flow
hollow-vaned impellers and integral diffusers increase propellant
flow-rate and pressures. Stationary diffuser vanes inside the pump
provide uniform pressure distribution, reduction of fluid velocity
around the impellers, and reduction of fluid turbulence in the pump
volutes. Balance ribs on the inboard side of the impellers hydrauli-
cally balance the axial thrust on the pump shaft. During operation
the fuel additive lubricates and cools the turbopump gears and
bearings. The lox and fuel single-entry, centrifugal pumps mount
back-to-back, one on each side of the gearbox. Bolts secure the
fuel pump to the gearbox while radially-inserted steel pins secure
the lox pump to the gearbox. The steel pins permit the lox pump
housing to expand and contract during extreme temperature
changes without distortion or misalignment. To prevent formation
of an explosive environment, four lines from the turbopump drain
expended lubricant and fuel and lox leakage overboard. Two lube
drain ports manifold into a single line that incorporates a lube
drain relief valve. This valve also maintains the gearbox pressure
between 2 and l0 psig. Two lox lines drain any lox leakage from
4-18
ITEM
CHAMBER AREA (INJECTOR END)
CHAMBER DIAMETER (INJECTOR
END)
THROAT AREA, At
THROAT DIAMETER
EXIT AREA
EXIT DIAMETER
NOZZLE AREA EXPANSION
RATIO
OVERALL LENGTH
TUBE WALL THICKNESS
NUMBER OF TUBES
CHARACTERISTIC
V
LENGTH L = c
At
COMBUSTION CHAMBER
VOLUME V c
DESCRIPTION
332 in. 2
20.56 in.
204.35 in2
16.13 in.
1634.3 in2
45.62 in.
8:1
86.15 ;n.
0.012 in.
292
39.10 in.
4.62 ft3
Figure 4-20
CH-14221-1
i I :l:|U_llldl/_1,',l:] | :i[o] g | :1:| I1_ [
PARAMETER
SEAL LEVEL THRUST
SEA LEVEL SPECIFIC IMPULSE
TOTAL PROPELLANT FLOWRATE
MIXTURE RATIO
LOX FLOWRATE
FUEL FLOWRATE
INJECTOR END CHAMBER
PRESS.
NOZZLE STAGNATION
PRESS.
CHARACTERISTIC
VELOCITY (C*)
C* INJECTOR END
PRESS.
C* NOZZLE STAGNATION
PRESS.
C t EFFICIENCY (NOZZLE)
THRUST COEFFICIENT (Cf)
Cf INJECTOR ENDPRESS.
Cf NOZZLE STAGNATIONPRESS.
Cf EFFICIENCY (NOZZLE)RATIO OF INJECTOR END
PRESS. TO NOZZLE
STAGNATION PRESS.
JACKET PRESSURE DROP:
AT 225 Ib/sec
AT 227.4 Ib/sec
I!11: f:l :lf:(q 1:1:! i.11 [¢ll
RATING
204,300 Ibf
268.9 sac
760.3 tb/sec
2.342 : 1 (O/F)
532.8 Ib/sac
227.5 Ib/sec
701.8 _la
652.5 psia
6,069 if/sac
5,647 if/see
97.20 %
1.425
1.532
101.4%
1.080
135.0 p_ig
138.0 pslg
Figure 4-21
CH-14222-2
the lox cavity downstream from the primary lox seal. One line
drains any fuel and lubricant leakage past the lube seal into the
fuel drain manifold, which dumps the fuel overboard through asingle line.
Gearbox. The gearbox contains the gear train that provides the
turbine differential speed to drive the main pump shaft (figure
4-22). All gears are of full depth configuration. The intermediate
and pinion gears contain inner races for the roller bearings. The
main pump shaft utilizes a roller bearing and a ball bearing. The
ball bearing restricts axial movement of the shaft. Lube passages
cast within the gearbox wails direct the fuel and Oronite mixture
through jets onto the bearings and gears. The jets apply lubricant
to the disengaging side of the gears to prevent hydraulic lock. An
accessory drive pinion, integral with the intermediate gear, drives
the two counter-rotating accessory drive gears. The lower accessory
Section IV S-IB Stage
H-I ENGINETURBOPUMP
/
GAS
FIRST STAGE
TURBINE WHEEL
TURBINE
SECOND STAGE
TURBINE WHEEL_
TURBINE
STAGE SEA
TURBINE
SECOND STAGE
NOZZLE
FIRST STAGE
NOZZLE
HAFT SEALS
8 BEARING
SEALS
NO. 7 BEARING
BEARING
NO. 5 BEARING
A-A
ACCESSORY DR}VE
ADAPTER (2 PLACES)
NO. 4 BEARING
3 BEARING
iMPELLER
INDUCER
FUEL SEAL
FUEL PUMP
NO.
TURBOPUMP CHARACTERISTICS
RATING AT 205KITEM
Inlet Density
Inlet Press.
(total)
Discharge Density
Discharge Press.(total)
Shaft Speed
Developed PumpHead
Volume Flow
Flowrate
EfficiencyShaft Power
NPSH Required
OXIDIZER FUEL
70,79 50.45 Ib/cu fl
65.0 57.0 psia
70.46 50.55 Ib/cu f_
970.0 1,012.0 psla
6,680 6,6_) rpm
1,851 2,719 ft
3,458 2, 149 gprn
545.4 241.6 fb/sec
77.88 71.78%
_:_, 357 1,664 flhp35.0 _ 35 40
Steady state ohly. EngTne starting requires 80 psla.
Starting and steady state requirement.
VOLUTE
SECTION A-A
VOLUTE
SEAL
PUMP
TURBINE CHARACTERISTICS
ITEM RATING AT 205K
tntet Press. (total)
Exit Press. (talc)Press. Ratio
Total Inlet/StatlcExhaust
inlet Press. Static
Shaft Power
Efficiency
Speed
624 psla
35.2 psTa
17.74 psla
540.7 psia
4,141 hp69.77 %
32, 632 RPM
CH.14199-2
Figure 4-22 (Sheet 1 of 2) 4-19
Section IV S-IB Stage
H-1 ENGINETURBOPUMP
SEAL_
NO.8.AR'NG--_r_TURBINE 1_ _
,.APT_
SPACER----_
Y_BEARINGHOUSING
AND
B-B "_@'/r--_/
NO. 5 BEARING,-_
PINION GEAR_ .NO. 6 BEARtNG_
NO. 8 NO. 7
BEAR,NOHOOS,NG/ /-X
FUEL
PUMp
MAIN GEAR_
NO. 2 BEARING
I1
Dm
====:=:
W
SECTION 8-B
'_ FUEL
VOLUTE
ASSY.
S ACCESSORY DRIVE
GEARACCESSORY DRIVE
COVER ACCESSORY DRIVECOVER I_EMOVED
/-;c_%_....._ _ / ......._ACCESSORY DRIVE
GEARS (ADG)
VIEW C
_NOo 3 BEARING
f NO. 1 BEARING
,_ LOXPUMP
--r__ _'_ MAIN SHAFT J GEAR RATIOS
RPM = N
% NC, 8, OR ADP : 2.303 N D
N A = 4.885 N D
l NAD G = 0.609 N D
GEARBOX
HOUSING
C-H 14200
4-20 Figure 4-22 (Sheet 2 of 2)
Section IV S-IB Stage
on the downstream side of the gate housing supplies fuel pressure
for thrust OK pressure switches actuation.
Igniter Fuel Valve. The normally closed igniter fuel valve (IFV)
attached to and operated mechanically by the main lox valve begins
to open when the main lox valve opens to 50 deg and is fully
open when the lox valve opens to 70 deg. Control-pressure fuel
flows through the IFV to the hypergol container and to the ignition
monitor valve for main fuel valve opening. Spring pressure closes
the igniter fuel valve at engine cutoff when the main lox valvecloses.
A drain line prevents any control-pressure fuel leakage from past
the IFV seals from entering the MLV linkage housing. The line
drains the leakage through the fuel drain manifold into the engineexhaust stream.
Gas Generator and Control System.
The gas generator and control system controls engine start sequenc-
ing and supplies power to drive the turbopump.
Solid Propellant Gas Generator. A solid propellant gag generator
(SPGG) installed on the liquid propellant gas generator (LPGG)
combustor of each engine, starts the engine by supplying largevolumes of gas to spin the turbine. Upon receipt of the engine
ignition command, two initiators installed in the SPGG ignite
pellets in an igniter inside the SPGG and the pellets in turn start
the grain propellant burning. As the grain burns, SPGG internal
pressure increases until a diaphragm bursts at 600 to 700 psig and
releases the gases through an orifice to the turbine. The turbine
drives the turbopumps, through a gear train, and liquid propellants
begin flowing in the engine. The SPGG continues to burn for
approximately 100-200 msec after bootstrap fuel and lox enter the
liquid propellant gas generator, thus igniting the liquid propellants.
SPGG Initiator. The SPGG initiator is an electrically activated
pyrotechnic squib that ignites the SPGG. A 500-Vac, 2-A ignition
command impulse applied to a bridgewire in the initiator ignites
the pyrotechnic charge. To ensure against inadvertent firing of the
initiator, a cold cathode trigger diode in the initiator circuit prevents
current through the bridgewire until the voltage exceeds 250 Vac.
A 100 ohm resistor in each electrical leg protects the bridgewire
from high level static electric discharge. Functioning time is 15
msec (max). The pyrotechnic material and bridgewire firing circuit
components are housed in a case made of 1018 steel bar material.
The housing has two sets of threads, one for installing the initiator
in the SPGG and one for attaching electrical cables.
Liquid Propellant Gas Generator. The liquid propellant gas genera-
tor produces combustion gases during steady-state operation to
drive the two-stage turbine, which supplies power through a gearreduction train to drive the propellant pumps. The LPGG consists
of a gas generator control valve, injector assembly, a combustor,
and two squibless igniters. Propellants entering the LPGG are
ignited by hot gases produced by the solid propellant gas generator
and squibless igniters during engine start. The hot gases from the
SPGG ignite the squibless igniters prior to liquid propellant entry
into the LPGG. The igniters burn for 2.5 to 3 sec to ensure liquid
propellant ignition. See figure 4-24 for LPGG characteristics.
Gas Generator Control Valve. The normally closed gas generator
control valve contains two poppets that admit fuel and lox bootstrap
propellants into the gas generator combustor during engine opera-
tion. TC combustion pressure actuates the control valve by reposi-
tioning a piston that opens the fuel poppet first. A yoke integral
with the piston opens the lox poppet. Fuel poppet cracking pressure
is 105±20 psig, lox poppet cracking pressure is 200--+20 psig, and
fully-open operating pressure for the control valve is.275-+25 psig.
Bootstrap fuel flowrate is 13.52 lbm/sec while bootstrap lox flowrate
is 4.61 lbm/sec. A bellows assembly enclosing the lox poppet stem
PGG CHARACTERISTICS
4-22
pARAMETER
TOTAL FLOWRATE
MIXTURE RATIO
GG COMBUSTION PRESS.
INJECTOR END
GG COMBUSTOR TEMP
FLOWRATE:
LOX
FUEL
PRESS DROP ACROSS
INJECTOR:
LOX
FUEL
RATING
18.13 Ib/sec
0.341 (0/0
646.3 psia
1,198OF
4.61 Ib/sec
13.52 Ib/se c
93 psid
90 psid
CH.14220-2
Figure 4-24
and closure spring, and seals on the actuator piston, prevent leakage
of fuel and lox into the control valve actuator. A drain line ports
any fuel leakage into the valve actuator to the fuel drain manifold
where it is dumped overboard into the engine exhaust stream. The
control valve design ensures a fuel-rich cutoff to prevent excessive
temperature buildup in the combustor resulting in turbine burning.
Spring pressure closes the control valve at engine cutoff when the
thrust chamber pressure decays.
Gas Generator Injector. Fuel and lox from the gas generator control
valve enter the injector and flow through passages that provide
a uniform mixture-ratio of 0.341 (lox/fuel). The injector cavity
design permits an oxidizer lead into the combustor during start
to prevent detonation. From the injector, two fuel streams impingeon a single lox stream. The injector uses 44 impingement points.
Fuel entering the combustor through 36 holes around the periphery
of the impingements provides film coolant for the injector. During
countdown the GG fox injector receives an ambient GN 2 purge
to prevent entrance of SPGG contaminants into the injector. See
Engine Purges.
Combustor. The bootstrap propellants burn in the combustor and
exit to the gas turbine. Two squibless igniters installed in thecombustor just below the GG injector assure propellant ignition
during start. The combustor is a welded assembly with flangesfor installation of the SPGG and for attachment to the gas turbine.
Operating temperature is 1198-+16 ° F and operating pressure is
646.3-+4.9 psia.
Squibless Igniters. Two squibless igniters installed in the injector
mounting flange on the combustor burn for 2.5 to 3 sec after their
ignition by the solid propellant gas generator. They ensure ignition
of fuel and lox if the SPGG burn should have expired before
bootstrap propellant entry into the combustor. A 2-A link wire
provides a monitoring capability for engine premature ignition.
The sixteen igniter circuits (two per engine) are series connected
to the ML. If any of the link wires break or burn through, for
any reason, a PREMATURE IGNITION lamp on the S-IB firing
panel will illuminate and the PREMATURE IGNITION SAFE
lamp on the S-IB stage networks panel will extinguish. An emer-
gency cutoff command will be initiated automatically when the
PREMATURE IGNITION switch is in the ARM position. Thelink wire has no function in the ignition of the squibless igniter.
A 7.5- f_ resistor in the series link wire circuit removes the possibility
of igniting an igniter by the monitor circuit power.
Ignition Monitor Valve. The ignition monitor valve (IMV) opens
t.he main fuel valve when ignition has been achieved in the combus-tion chamber. The IMV mounts below the MFV actuator and
inter'faces with the MFV opening port by an adapter. The three-way,
normally closed valve has four ports; control-pressure fuel inlet
(from IFV) and outlet (to MFV actuator), combustion chamber
pressure sensing port, and a drain port. When combustion chamber
pressurereaches15+-0.5psig, the valve shuttles, closing off the drain
port, and permits control-pressure fuel flow into the MFV actuator.
The IMV remains open throughout H-1 engine operation. During
engine shutdown the valve closes under spring pressure. The drain
port opens and dumps the MFV opening actuator fuel overboard
through a drain line permitting the MFV to close.
Thrust OK Pressure Switches. Three normally open, two-position
pressure switches on each engine, sense fuel pressure downstream
of the main fuel valve. Each hermetically sealed pressure switch
contains a single-pole, double-throw switch with positive actuated
snap-action electrical contacts. A checkout port enables CALIPS
checkout testing without disconnecting engine system lines or pres-
surizing the fuel inlet manifold. The electrical outputs of the
switches indicate satisfactory fuel pressure as a component of satis-
factory engine thrust. The switches actuate when fuel pressure
downstream from the main fuel value reaches 800±15 psia (indicat-
ing approximately 90 percent engine thrust attained). Switch deac-tuation occurs between 45 psi (max) and 25 psi (min) below the
actual actuation pressure. During checkout, pressure applied to theswitches from the CALIPS console actuates the switches at 800-+-45
psia for fast actuation and 800_30 psia for second and third actua-
tion. Deactuation pressures during checkout are 15 to 65 psi below
actuation pressure for the first cycle and 20 to 45 psi for second
and third cycle. During the ignition sequence operation (T-3 toT-0) the IU EDS distributor monitors the thrust OK pressure
switches for thrust buildup. Voting logic for each three-pressure-
switch-group on each engine determines when all engines have
attained approximately 90 percent of rated thrust and providesan ALL ENGINES RUNNING signal to the program distributor
in the ML. Tile logic circuits remove power from the L/V engine
indicator lamps on the command module main display console,
indicating that the engines are running and have attained approxi-
mately 90-percent thrust level. If all engines are running at time
for commit, T-0, the COMMIT signal will command the launchvehicle release circuits. If, however, the ALL ENGINES RUN-
NING signal is not present at time for commit, cutoff will occur
automatically. ALL ENGINES RUNNING indications are moni-
tored on the S-IB firing panel. The THRUST OK pressure switches
provide discrete inputs to the ground computer systems and the
telemetry system.
Hypergol Container. The .hypergol container is an integral part of
the thrust chamber injector (figure 4-18). The cylindrical housing
accommodates a 6-in? hypergol cartridge of triethylalumin, a (fig-
ure 4-25) and a HYPERGOL INSTALLED detector switch. O-ring
seals on the cartridge prevent leakage during engine operation.
Two diaphragms, one at the inlet and one at the outlet of the
cartridges, contain the hypergol until fuel pressure, 300-+-25 psig,
burst them during engine start. Hypergol flows, under turbopump
fuel pressure, through the igniter fuel manifold to seven passages
that direct hypergol to igniter fuel ports in the injector. Ignition
occurs when the hypergol contacts lox. Fuel flows through the
hypergol container and igniter fuel ports during engine operation.
Insertion of the cartridge into the container actuates the detector
switch, which provides an output to the S-IB firing preparation
panel, HYPERGOL INSTALLED indicator lamps illuminate in-
CH.14219
Figure 4-25
Section IV S-IB Stage
dicating all eight engines have received the hypergol cartridges.
The switches also provide inputs to the digital event evaluator.
Conax Valve. The two-way, normally closed Conax valve effects
engine cutoff by directing control-pressure fuel to the main lox
valve closing actuator (figure 4-16). Two pyrotechnic actuators
installed into a single body connect the inlet and outlet ports by
driving a ram through metal membranes separating the ports.
Actuation of either or both pyrotechnic actuator assemblies will
allow fuel flow through the valve. Position indicators installed into
the valve body opposite the actuators provide monitoring of the
valve position. When the actuator fires, the ram strikes a plunger,which in turn breaks a link wire in the indicator. All sixteen position
indicators (two per engine) are connected in series to ground moni-
toring circuits. With all link wires intact, no ground indication will
be monitored. However, if any link wire breaks, a red ANY CONAX
FIRED lamp will illuminate on the S-IB firing panel in the LCC.
As a safety precaution, no electrical connections exist between the
position indicator and the pyrotechnic actuator. To prevent inad-
vertant application of power or stray voltage from firing the Conax
valve, a ground "Conax valve safe" command, constituted by
deenergized emergency cutoff and ignition command relays, ener-
gizes four vehicle relays thatground actuator firing circuits. Either
the T-5 sec signal or an emergency cutoff command disables the
ground "Conax valve safe" command. In the vehicle, normallyclosed contacts of the command engine cutoff relays also keep the
actuator firing circuits at ground potential until the guidance com-
puter or RSO commands engine cutoff or if the thrust ok pressureswitches on an engine initiate cutoff. The command engine cutoff
relays reposition their contacts and apply power to the Conax valve
actuator. Bridgewires detonate the actuator pyrotechnics. The valves
open, and control-pressure fuel flows to the main lox valves closingactuators.
Exhaust System.
The H-1 engine exhaust system (figure 4-9) ducts the fuel-rich
turbine exhaust gases overboard into the thrust chamber exit flowstream. A welded stainless steel turbine exhaust hood ducts the
gases into a heat exchanger. A bellows section with an integral
liner in the turbine exhaust hood permits movement of the system
due to heating. The heat exchanger, also a welded stainless steel
shell, houses a helix-wound four-coil system. Exhaust gases heat
the coils. Lox, under turbopump pressure, flows through a unitized
check valve into three of the coils and is converted to gox for
lox tank inflight pressurization. An orifice in each of three coilinlets controls the lox flowrate. The fourth coil is not used. Exhaust
gases exit the heat exchanger into a turbine exhaust duct on inboard
engines, or an aspirator on outboard engines. The curved stainlesssteel turbine exhaust duct directs the exhaust gases into the thrust
chamber exit flow stream of the inboard engines. The aspirator,
a welded Hastelloy C-shell assembly installed on the periphery
o e the outboard engine nozzle, extends below the thrust chamber
exit. The forward end of the aspirator is welded to a channel band
approximately 20 in. forward of the fuel return manifold. The aft
end of the aspirator is not secured. A 0.440-in. clearance between
the fuel return, manifold and the aspirator permits the turbine
exhaust gases to escape into the thrust chamber exit flow stream.
Lubrication System.
The fuel additive blender unit (FABU) eliminates the requirement
for a lube-oil tank, pressurization, equipment, plumbing, and con-
trols. Extreme-pressure additive, Rocketdyne ST0140RB0013,
blended with RP-I fuel lubricates and cools turbopump gearbox
components (figure 4-22). A thermostatically controlled 300-+-30 W
heater maintains the correct viscosity of the additive by controlling
the temperature between 12___4° F and 130---+4° F. GSE supplies
115V, 60Hz power to the heater for preflight additive conditioning.
4-23
Section IV S-IB Stage
FUEL ADDITIVE HEATER POWER ON lamp on the AC module
and FUEL ADDITIVE HEATERS lamp on S-IB networks panel
illuminate at heater power application. A current sensor in each
heater circuit provides an input to DEE-6 that the heater is opera-
tive. The ALL ENGINES RUNNING signal removes heater power
just before the umbilicals disconnect. Lubricant flow begins when
control-pressure fuel increases to 70 to 150 psig. Fuel flows through
a 35-mesh inlet strainer made of Monel and pressurizes the additive,
and aligns a spool containing a metering orifice with the additive
outlet. The additive then flows through a 100-mesh Monel outlet
screen and blends with the fuel (2.75_--+-0.75 percent by volume).
The mixture leaves the FABU and enters the turbopump througha 40 (nom) to 75 (max) micron-mesh filter. The filter element is
made of 18-8 CRES stainless steel. Lubricant consumption rate
is 5 to 6 gpm. After injection into the turbopump gearbox, the
lubricant drains overboard through the lube-drain relief valve and
drain lines that extend down the engine thrust chamber exterior
and dump into the engine exhaust stream. At engine cutoff the
decaying fuel control pressure permits spring closure of the FABU,
which shuts off lubricant flow to the gearbox.
Electrical System.
The H-1 engine electrical system consists of armored and unar-
mored electrical harnesses that interface with engine components
and stage and ESE circuitry. Harnesses that have flight as well
as preflight functions are armored to prevent damage to the con-
ductors that may compromise engine operation or cause failure
of the mission. Harnesses that have preflight functions only do
not have armor. Those preflight functions are heater operations
(FABU, turbopump bearing no. 1, and MLV actuator), auxiliary
hydraulic pump operation, Conax valve position indications, squib-
less igniter link monitoring (premature ignition), hypergol cartridgesinstalled indications, and start commands to the SPGG initiators.
The armored flight harnesses transmit thrust ok signals, hydraulic
servoactuator commands and feedback, and cutoff commands to
the Conax valves.
H-! ENGINE ANALOG MI:ASURMENTS
H-1 Engine Measuring.
The H-I engine measuring systems monitor 13 conditions on each
outboard engine and 11 conditions on each inboard engine. This
information is telemetered to ground receiving stations through
the S-IB stage PCM/DDAS assembly and RF assembly PI during
flight. During checkout, telemetry is received by coaxial cable from
the PCM/DDAS assembly. Each measurement has a number com-
prised of a letter representing a parameter, measurement number
within the parameter, and a dash number indicating the stage unit
location (Example: CI-I). Three parameters of engine measure-
ments are temperature, C; pressure, D; and RPM, T. Engine unit
numbers are 1 through 8 respective to engine location. See figure
4-26 for measurement numbers, titles, and other pertinent informa-
tion and figure 4-9 for measurement locations. Output signals frommeasurement transducers XCI, C9, XC54, XC89, DI, and TI2 are
routed through measuring racks 9A516 for engines 1 and 5, 9A520
for engines 2 and 6, 9A526 for engines 3 and 7, and 9A530 for
engines 4 and 8. The measuring racks contain signal conditioning
modules for each measurement input. The modules assure compati-
bility of the measurement signal with the TM multiplexer input
requirement. After signal conditioning, measurement signals XCI,
C9, XC54, XC89, and DI are multiplexed by TM multiplexer
13A484. Measurement signals T12 are applied directly to TM
assembly F1. Measurement signals DI2, DI3, and DI4 do not
require signal conditioning and are applied directly to TM mul-
tiplexer 13A440. Measurement signals D20, D34, and D35 do not
require signal conditioning and are applied directly to TM mul-
tiplexer 13A484. Measurement signals XCI, C9, XC54, DI, and
TI2 require remote automatic calibration system (RACS) checkout.
This permits checking the individual circuits for respons e and
4-24
NUMBER
XCI-1 THRU -8
C9-1 THRU -S
XC54-1 THRU -8
XC89-1 THRU -8
C540-1 THRU -8
CS41-1 THRU -8
C542-1 THRU -8
DI-1 THRU -8
D12-1 THRU -4
D13-1 THRU -4
D14-1 THRU -8
D20-I THRU -8
D34-I THRU -8
XD35- I THRU -8
D53-2 AND -6
E513-2 AND -6
T12-1 THRU -8
NAME
TEMP, LOX PUMP BEARING I
TEMP, GAS GEN CHAMBER
TEMP, LOX PUMP INLET
TEMP, GEAR CASE
TEMP, LOX SEAL DRAIN LINE 1
TEMP, LOX SEAL DRAIN LINE 2
TEMP, LOX SEAL DRAIN LINE 3
PRESS, COMBUSTION CHAMBER
PRESS, FUEL PUMP INLET
PRESS, LOX'PUMP INLET
PRESS, TURBINE INLET
PRESS, GEAR CASE
PRESS, GG FUEL INJECTOR
PRESS, GG LOX INJECTOR
PRESS, LOX PUMP INLET
VIBRATION, ENG THRUST BLOCK, LONG
TURBINE RPM
RANGE
-20 TO 200°C
0 TO 1000°(3
-185 TO - 1680C
0 TO 150°(:
:- 185 TO - 12°C
- 185 TO - 12aC
- 185 TO - 12°C
0 TO 800 PSIA
0 TO 100 PSIA
0 TO 150 PSIA
0 TO 800 PSIA
0 TO 200 ESIA
0 TO 900 PSIA
0 TO 900 PSIA
0 TO 150 PISA
-ITO +5G
0 TO 45K RPM
[_D "X" AUXILIARY DISPLAYPREFIX INDICATES
CH-14323-1
Figure 4-26
accuracy of signal transmission. The RACS Control panel operator
selects HI, LO, or RUN reference signals that correspond to pre-
dicted transducer signals and applies the signal to an individual
circuit. Upon receipt of instructions from the RACS, the S-IB stage
measuring rack selector (9A546) addresses the particular measuringrack and module to be checked. With the reference signal transmit-
ted to the signal conditioning module, the module returns a signal
corresponding to the predicted transducer output. This signal veri-
fies calibration of the individual circuit. Two H-1 engine parametersare considered critical and have redline values which, if exceeded,
will produce unsafe or unsatisfactory operations. Measurements
XC89-1 through XC89-8, extreme-pressure additive temperature
at each engine, must be within a 105 ° F to 160 ° F range. If the
lube additive is outside this range, the additive consistency will
cause improper mixture of fuel and additive resulting in improper
turbopump lubrication. Measurements XCI-1 through XCI-8, tur-
bopump bearing no. 1 temperature at each engine, must have a
minimum temperature of 40 ° F at T-3 min. Redline value at
ignition is 0 ° F, which is based on heat loss (assuming heater failure)
at T-3 min. The position of each thrust ok pressure switch on
each engine is telemetered back to the ground receiving station
as event measurements. See figure 4-27. In addition to being re-
corded, these signals are monitored by ENG THRUST OK indica-
tors, one for each pressure switch, on the EDS monitor panel.
Inputs to the EDS monitor panel are received through DDAS.Prime interest time of these measurements is from liftoff until S-IB
stage outboard engine cutoff. These measurements are fed directly
from the thrust OK pressure switch to the Remote Digital Submul-tiplexer 9A700 and are telemetered through the PCM/DDAS as-
sembly and RF assembly PI to the ground stations. During check-
out, thrust OK pressure switch positions are also monitored, throughhard-wired connections, on ENG THRUST OK indicators on the
EDS Preparation Panel. This panel also has an indicator for each
H-1 ENGINE EVENT MEASUREMENTS
MEAS. NO.
VK138-I
VK139-I
VK140-2
VK141-2
VK142-3
VK|43-3
VK144-4
VK145-4
VK146-5
VK147-5
TITLE
ENG 1, SW 1 THRUST OK
ENG 1, SW 2 THRUST OK
ENG 2, SW 1 THRUST OK
ENG 2, SW 2 THRUST OK
ENG 3, SW 1 THRUST OK
ENG 3, SW 2 THRUST OK
ENG 4, SW I THRUST OK
ENG 4, SW 2 THRUST OK
ENG 5, SW I THRUST OK
ENG 5, SW 2 THRUST OK
VKI48-6
VK149-6
VKIS0-7
VK151-7
VK152-8
VK|53-8
VK171-|
VK172-2
VK173-3
VK 174-4
VK175-5
VK176-6
VK177-7
VK178-8
ENG 6, SW 1 THRUST OK
ENG 6, SW 2 THRUST OK
ENG 7, SW 1 THRUST OK
ENG 7, SW 2 THRUST OK
ENG 8, SW 1 THRUST OK
ENG 8, SW 2 THRUST OK
ENG 1, SW 3 THRUST OK
ENG 2, SW 3 THRUST OK
ENG 3, SW 3 THRUST OK
ENG 4, SW 3 THRUST OK
ENG 5, SW 3 THRUST OK
ENG 6, SW 3 THRUST OK
ENG 7, SW 3 THRUST OK
ENG 8, SW 3 THRUST OK
C.H 14324
Figure 4-27
pressure switch on each engine. Measurements D 1-1 through DI-8and VK138 through VK153 are flight control measurements that
are monitored in real time at Mission Control Center in Houston.
The three lox seal drain line temperature measurements on each
engine (C540, C541, and C542) are interlocked in the automatic
countdown sequence between engine ignition and liftoff. If anytwo of the three measurements on any engine indicate the presence
of lox (temperature below -250 ° F) in the drain cavity, all engines
will be cut off.
STATIC TEST.
S-IB-6 was static fired on June 23, 1966 (Test SA-36) for 35.58
sec duration and again on June 29, 1966 for 141.24 sec (Test SA-37).
Performance and engine operation was satisfactory on Test SA-36.
During Test SA-37, Engine 2 (H-7072) and Engine 4 (H-7075)
experienced step decreases in power level at 105 sec and 25 sec
respectively.
S-IB-6 is the first stage equipped with 205K engines. These engines
are essentially the same as the previous 200K stages except for
changes necessary to increase propellant flowrates and to allow
the higher operating level. Specifically, the turbopump lox and
fuel impellers were retrimmed for increased flowrates and the gas
generator injector differential pressures were lowered.
Test SA-36.
This test was conducted at the Static Test Tower East (STTE) at
MSFC, Huntsville, Alabama. Cutoff was initiated by the firing
panel operator as scheduled. Duration from ignition command toInboard Engine Cutoff (IECO) was 35.46 sec and to Outboard
Engine Cutoff (OECO) was 35.58 sec. All engines performed withinthe 205,000+3000 lbf range (sea level reference conditions). No
recalibration was necessary. Engine 6 (H-4069) operated at 207.4K
sea level thrust, 2.8K higher than the Rocketdyne level.
Test SA-37.
This long duration test was also conducted on STTE. Time basetwo was initiated by uncovering of the low level sensor in Tank
02 at 135.38 sec after ignition command. At 3.2 sec after uncovering,
Section IV S-IB Stage
IECO was commanded by the switch selector. OECO was from
lox depletion occurring 2.66 sec after IECO. Time from ignitioncommand and actual event times are not necessarily representative
of those expected for flight, but were altered to fit the static test
conditions.
As mentioned in the initial paragraph, engines 2 and 4 exhibited
step changes in thrust during SA-37. These shifts were noticed inthrust chamber combustion chamber pressure and amounted to
approximately one percent.
In addition to the in-run shifts, all engines operated at a steady
state power level lower than on Test SA-36. Exhaustive studieshave failed to detect the reason for this overall reduction. The
flight prediction has considered both the shift and the lower powerlevel.
The step decreases have been observed in Rockeydyne tests and
were found to be caused by changes in fuel flow distribution inthe thrust chamber coolant tubes. The flow distribution change
resulted in fuel system resistance increases. The addition of a bame
in the thrust chamber exit manifold successfully eliminated the
shifts. However, the installation of the baffle is not scheduled until
S-IB-13 engines.
See figure 4-28 for a summary of the 205K H-I engine test history.
TEST
PROGRAM ENGINES DURATION
TESTED (SECONDS)
1 2,765 37
ACCEPTANCE TEST 97 28,657 364
R&D TEST PROGRAMS 27 62,373 843
STAGE STATIC TESTS 64 12,708 184
TOTAL 106,503 1,428
Figure 4-28
LOX SYSTEM.
The five S-IB lox tanks receive lox from the facility storage system
through the fill-and-drain line storing it for consumption by the
eight H-1 engines during boost phase (figure 4-29).
The lox system tanks consist of four outer units (O-1, 0-2, 0-3,and 0-4) and a center unit (O-C), with a nominal system capacity
of 66,277 gal and a minimum ullage volume of 1.5 percent. Suf-
ficient ullage pressure is provided to ensure structural integrity ofthe lox tanks, and to maintain a net positive suction head of 35
ft at the lox pump engine inlet. The skin-milled, butt-welded alumi-
num alloy segments of the tank walls vary in thickness from topto bottom in relation to stress concentrations. The tank bulkheads
are hemispherical, with skirts forward and aft providing space for
pressurization and vent manifolds on the forward end. The aftskirts accommodate the sumps and interconnecting manifolds. Each
of the outer tanks supplies lox to one inboard and one outboard
engine. The center tank holds approximately 35 percent of thelox and is 105 in. in diameter by 678 in. long. The overall length
is 750 in. Clustered around the center tank are four tanks, each
having a capacity for 16 percent of the lox requirements. The outertank dimensions are 70 in. in diameter by 678 in. long, between
4-25
SectionIV S-IB Stage
S-IB LOX SYSTEMDIAGRAM
LOX VENT
VALVE(,_--_ _ JI _ II LOXSENSING
BAFFLES
II 'tl II
_._. _ il_ul pANL_NDGJ II _ _ SA_RTJEVNORTEi! _O N N ECT.._ _"_ // ,IIf i_)EXALvE II II
PRESSURE SYSTEM
[_ SEE S-IS LOX PRESSURIZATION SYSTEMCH.14640.1
Figure 4-29
bulkheads, and an overall length with skirts of 747 in. Each tank
has a capacity of 10,821 gal, but 1.5 percent of the capacity is
reserved for ullage.
Lox Fill.
Initial Fill. The S-IB lox loading is an automatic operation con-
trolled by the lox tanking computer. During the initial fill, the
stage vent valves, the main fill valve, the slow fill valve and the
pump discharge valve are opened to allow the lox to flow from
facility storage to the stage tanks. As the lox is loaded, it is distrib-
uted equally to each tank by the manifold. The lox temperature,
as it is loaded, remains almost constant for the entire fill operation
between -285 and -297 ° F. The flowrate for this precool opera-
tion is approximately 500 gpm, for approximately 30 min. until
22-percent loading is obtained. The propellant tanking computer
system (PTCS) generates a signal that terminates the precool opera-
tion and starts the main fill operation.
Main Fill. When the main fill valve, the pump discharge valve,
and the S-IB fill-and-drain valve are opened the S-IB umbilicalvent line drain valve is closed, the flowrate is increased to 14,250
lbm/min. This rate decreases when 95-percent load is reached in
approximately 15 min. The topping, the last 5 percent of the634,125-1bm load, is slow-fill loaded at a rate of 5260 lbm/min
and at a pressure of 50 psig. This sequence is accomplished by
closing the main fill valve. The 99-percent level signal causes the
slow fill valve to close and the fill sequence is complete. The lox
level is now maintained to the 100-percent level until 3 min and7 sec before liftoff.
Lox Bubbling. Lox bubbling is initiated approximately 153 secbefore liftoff and must not be terminated more than 100 sec before
ignition. Helium is the medium for the bubbling action that is
introduced into the lox suction lines at a ground regulated pressure
of 225 psig, at ambient temperature and a flowrate of 3.263 lb/min.
The helium gas flows upward through eight separate branch lines,
each containing a metering orifice, and then into the pump inletat the end of each suction line. This helium flow maintains sub-
cooled lox at the turbopump inlets to prevent pump cavitation
during engine start. The helium bubbles then rise through the
suction lines and normally open lox prevalves and through the
lox tank where it contributes to the ullage pressure.
Lox Tank Pressurization.
Prepressurization. The S-[B tank is automatically sequenced for
pressurization after the lox bubbling from valve panel 2, while
the stage is on the ground, lnflight pressurization is provided by
gox converted from lox in heat exchangers located on the H-I
engines (figure 4-30). The prepressurization tank pressure is 55.3
to 57.7 psia at ambient temperature; the time required is 0.834
min. The pressurization is initiated by closing the lox vent valvesand the lox vent-and-relief valve. Ground helium flows to the center
lox tank and'from there to the outer lox tanks through the upper
interconnect lines and manifold. The tank pressure is controlled
by the lox prepressurization switch between 55.3 and 58.5 psia.
If the prepressurization switch fails, the lox vent and relief valve
will mechanically open between 60 and 62.5 psia. The ground lox
vent pressure switch will actuate at 67.5 psia and allow pneumatic
4-26
Section IV S-IB Stage
S-IB LOX PRESSURIZATION _.ND VENT ;YSTEM DIAGRAM
VENT VALVES NO. 1
AND NO. 3 GROUND
LOX VENT & _L PRESSURE Q-D
RELIEF VLV
OPENING
CONTROL J
Q-D_ ___1 /._ SOLENOID
SENSING _ VALVELINE
/----VENT VALVES NO. 2 _ _T"
/AND NO. 4 GROUND l = • _ _ _ LOX PRE-PRESSURIZATION JJ
JCONTROLPRESSUREO-OII LO×VENTAND_ _ II /SW,TC. IIJl RELIEF VALVE _ _ J LOXEME _TGcEHNCY JJ
SENSING VALVE LOX TANK TO H-1 ENGINE SYSTEM
Q*D
VALVE
tFROM H-I ENGINE SYSTEM
(TYP 8 pLACES) CH-14638-1
Figure 4-30
pressure fromthe ground GN 2 system to open the lox vent-and-relief valve.
Flight Pressurization. During the boost phase, lox is supplied to
the engine heat exchangers. Gox from each heat exchanger flowsinto a common manifold, through the gox flow control valve, andinto lox tank O-C. The outer lox tanks receive gox through the
upper interconnect lines and manifold. The gox flow control regula-tor controls the flow and maintains a pressure of 50 psia. Overpres-
surization is prevented by the lox vent-and-relief valve.
Measurements.
Figure 4-31 lists the lox system flight measurements and indicates
the information that is displayed.
Lox Characteristics.
Figure 4-32 lists the physical and chemical properties of lox, and
figure 4-33 shows the oxygen vapor pressure curve.
Lax System Components.
Lox Fill and Drain Valve. The ball-rotor gate valve is located in
the fill-and-drain line leading to the sump of lox tank 0-3. The
valve is spring-loaded to the closed position and is opened by
VK 16-.0,_
LS00-0C
LSO0-01
L500-03
It:>
NOMENCLATURE
GOX, LOX TANK
LOX LEVEL CUTOFF IND NO. 2-
LOX LEVEL CUTOFF IND NO. 3
LOX LEVEL PROBE
LOX LEVEL PROBE
LOX LEVEL PROBE
AUXILIARY AND MCC-H DISPLAY
ESE DISPLAY
Figure 4-31
RANGE DISPLAY J
_ +100 PSIA [_ I
ON OFF [_ I
ON O_F _> ]
o/+5VDC
0,/+5 VDC
0/+5 VDC
CH-14533-1
pneumatic pressure applied to the actuator assembly. The actuator
is designed for an operating pressure of 750 psig, a proof pressure
of 1125 psig, and a burst pressure of 1875 psig. The valve positionis indicated by an electrical switch for an LCC readout. The closing
4-27
Section IV SIB Stage
and opening response •time is 500 msec with a flow chamber pressure
of 100 psig. The valve is designed to handle lox at ambient temper-
atures from -100 to 125 ° F with a nominal operating pressure
of 150 psig, a proof pressure of 225 psig, and a burst pressureof 375 psig.
Lox Prevalve. The normally open, ball-rotor gate valves are located
next to the sumps in each of the eight lox feed lines. The primary
function is to stop the flow of lox to the engine and also provide
back up capability to the engines main lox valves. The eight pre-
valves are operated by a signal from the TOPS deactuation so
that prevalve closure is accomplished only after the engine hasstarted thrust decay. The valve is spring-loaded and driven to the
closed position by the pneumatic actuator; when the pressure is
released, the valve retums to the normally open position. Gaseous
nitrogen is the pneumatic medium operating at.a pressure of
775-+-25 psig. The actuator is designed for a proof pressure of 1125
psig and a burst pressure of 1875 psig. The actuator operating
temperature range is --65 to 125 ° F. A position indicator switch
monitors the fully closed and the fully open position for an LCC
readout. All prevalves must indicate OPEN to complete an interlock
for start of automatic launch sequence. The valve response time
for closing is 850 -+. 100 msec and for opening is 3,500 msec maxi-
mum under flow. The prevalve environmental temperature is - 100
to + 125 ° F at a nominal operating pressure of 150 psig, proof
pressure of 225 psig, and a burst pressure of 375 psig.
Gox Flow Control Valve. Gox is accumulated from eight engine
heat exchangers and its flow is regulated by the flow-control-valve
as it passes to the dome of lox tank O-C for distribution and equalpressurization tqt each of the five lox tanks. The control valve is
a modulating, spring-loaded, normally open, butterfly-type valve.
The butterfly is controlled by a pressure-operated bellows that tends
to close the butterfly. An aneroid sensor in the valve regulates
the pressure flow to the bellows from lox tank O-C. Thus an
increasing lox tank pressure tends to close the butterfly and a
decreasing pressure tends to open it, thereby increasing the gox
flow. A potentiometer monitors the position of the butterfly for
LCC readout. The proof pressure of the flow chamber and the
control pressure chamber is 750 psig. The bias pressure chamber
is proofed at 470 psig. The main power bellows has a negative
differential proof pressure of 305 psig and positive differential proof
pressure of 420 psig. The proof pressure of the pilot valve is 70
psia. The operating pressure for the flow chamber and the control
pressure chamber is 500 psig maximum. The bias pressure chamber
operates at a maximum pressure of 305 psig. The main power
bellows operate with a negative differential pressure of 202 psig,
and a positive differential pressure of 280 psig. The pilot valve
nominal operating pressure is 50 psia. The burst pressure of the
flow chamber and control pressure chamber is 1250 psig. The bias
pressure chamber burst pressure is 785 psig. Burst pressures of
the main power bellows are a negative differential pressure of 505
psig and a positive differential pressure of 700 psig. The gox flow
control valve operating temperature range is from 10 to 200 ° F.
Vent Valve. The four vent valves vent the pressure in the lox tanks
during lox loading and replenishing operation. The normally closed,
gate-type valves are spring-loaded and are opened by ground supply
GN2. A position indicator switch provides LCC readout for a fully
open or a fully closed gate. A 110-Vac thermostatically controlled
heater prevents the valve mechanism from freezing. The thermostat
energizes at a minimum of 70 ° F and deenergizes at 145 ° F. The
valve will open in 75 msec with a control pressure of 775-4-25
psig, and a minimum of 500 psig, while the flow chamber medium
is in a temperature range of -250 to 250 ° F. The control actuator
uses GN 2 and has a proof pressure of 1125 psig, with a burst
pressure of 1875 psig. The valve flow chamber has an operating
pressure of 60 psig, with a proof pressure of 90 psig and a burst
pressure of 120 psig.
PROPERTIES OF .OX.
OXYGENVAPORPRESSURE
4-28
COMMON NAME:
LOX, LIQUID OXYGEN
CHEMICAL FORMULA:
02
MOLECULAR WEIGHT:
32.0
PHYSICAL PROPERTIES:
FREEZING POINT ..................... -361.76°F
BOILING POINT ....................... -297.4OF
CRITICAL TEMPERATURE .............. -181.04°1:
CRITICAL PRESSURE ................... 736.47 PSIA
LIQUID DENSITY ....................... 9.54 LBS/GAL
APPEARANCE ........................... PALE BLUE,CLEAR LIQUID
ODOR ............................. : ...... NONE
LIQUID TO GAS RATIO .............. 1:862
CHEMICAL PROPERTIES:
STABLE AGAINST MECHANICAL SHOCK IN PURE FORM.
IMPACT SENSITIVE TO UNPREDICTABLE DEGREE IF CONTAMINATED,
ESPECIALLY WITH ORGANIC MATERIALS.
MIXED WITH GREASE, OILS, PETROLEUM DERIVATIVE FUELS, ALCOHOL,
ETC. , IT FORMS A HIGHLY IMPACT SENSITIVE GEL WHICH MAY BE
DETONATED BY SPARK OR FLAME AS WELL AS MECHANICAL SHOCK.
THE EXPLOSIVE POTENTIAL OF THIS GEL HAS BEEN SHOWN TO BE
APPROXIMATELY EQUIVALENT TO NITROGLYCERINE.
Figure 4-32
C-H 14538
750-
700-
650-
600-
550-
500-
450-
400-
350-
300-
25o-
200-
150-
50-
370 350
a
A
O-361.76 0.00 FREEZING POINT
J a -297.4 14.70 BOILING POINT
,_ / (1 ATM.)
a -181.04 736.47 CRITICAL POINTi r i t i i i i I i _ f i i i J i I
330 310 290 270 250 230 210 190 170
MINUS DEGREES FAHRENHEIT (°F) C.H 14S37
Figure 4-33
Vent and Relief Valve. The prime function of the vent-and-relief
valve is to prevent over-pressurization of the lox tank should the
pressurization system malfunction. The valve is a spring-loaded,
normally closed, pilot-operated, in-line poppet-type valve. The
valve vents the lox tank during the loading and replenishing opera-
tion and also relieves excessive pressure during preflight pressuriza-
tion and after liftoff. The valve is operated by 775±25 psig GN 2
from the ground system or by He and gox pressure through a
sensinglineconnectedbetweenthe valve pilot and the ullage area
of tank 0-3. The valve will crack at 60 psia minimum and will
reseat at 59 psia. A hermetically sealed position indicator switch
provides an LCC readout of the open and closed position of the
valve. The flow chamber and sensing chamber have a design
operating pressure of 63 psig, with a proof pressure of 95 psig
and a burst pressure of 160 psig. The vent-and-relief valve will
flow 16 lbm/sec (min) in the open position. The valve will move
from fully dosed to fully open position in 350 msec maximum
with 750 psig pressure on the actuator and with 58.5 psia in the
flow chamber. The maximum close cycle is 1.5 sec.
Pressure Transducer. Lox tank pressure is monitored at the prepres-
surization switch. The transducer is calibrated with ground source
GN 2 through the calibration valve. The pressure range of the
transducer is 0 to 100 psia, with a temperature range of -65 to
+ 200 ° F, and an electrical resistance of 5000 ohms.
Calibration Valve. The calibration valve is used to calibrate the
transducer. The nominal operating pressure is 3000 psig, with a
proof pressure of 4500 psig and a burst pressure of 7500 psig.
Section IV S-IB Stage
Operating temperatu[e range of the valve is -- 100 to 250 ° F.
Emergency Vent Switch. This pressure switch monitors only the
preflight pressurization of the lox tank. Should the ullage pressureexceed 67.5_1.5 psia because the lox vent-and-relief valve relief
function fails, the pressure switch will actuate a solenoid valve that
opens the vent-and-relief valve thereby relieving the excessive
pressure from the lox tank. The pressure switch deactuates at 63
psia to permit closing of the vent-and-relief valve. The switch is
calibrated with ground source GN 2. It has a proof pressure of 110
psia and a minimum burst pressure of 175 psia. The temperature
operating range of the switch is -65 to 165 ° F.
Prepresurrization Switch. The prepressurization switch controls the
ullage pressure of the lox tank, by actuating at 58.5 psia (max)
closing down the pressure and by deactuating at 55.3 psia (min),
permitting prepressurization to resume. This switch is disabled at
liftoff. The switch is calibrated with ground source GN 2. It has
a minimum proof pressure of 90 psig and a minimum burst pressure
of 150 psig. The operating temperature range of the switch is from-65 to 165 ° F.
;-IB FUEL ;YSTEM DIAGRAM
I
TO H-I ENGINE
SYSTEM (TYP.
8 PLACES)
3 7
_RIFICE (8_
FILTER
FUEL BUBBLING
LINE Q°D
Figure 4-34
CH-14637-1
4-29
Section IV S-IB Stage
Engine Cutoff Lox Sensor. The engine cutoff lox sensors, locatedin the bottom oflox tanks 0-2 and 0-4, initiate the inboard engine
cutoff when the lox is depleted to the level of the sensors. The
operating temperature range of the sensors is -65 to 165 ° F. The
linear accuracy of the sensors is plus 0.125 minus 2.00 in. of the
static set point repeatable to +0.125 in., when the level is decreasingat the rate of 5 in./sec. A heater and thermostat are incorporated
into the sensor. The heater operates during prelaunch only on 110V,
60 Hz, single-phase power supply and is rated at 50 W.
Lox Overfill Sensor. The lox overfill sensor is located in the top
of lox tank O-C, and actuates when the lox rises above the sensor
during lox filling, thereby terminating the filling sequence. The
operating temperature for the sensor is -65 to 165 ° F. The sensorhas a thermostatically-controlled heater as an integral part that
operates on 110V, single-phase power and is rated at 60 W.
FUEL SYSTEM.
The S-IB stage fuel system (figure 4-34) receives RP-I fuel from
the facility storage tanks, stores the fuel, and then supplies the
fuel to the eight H-1 engines. The RP-I fuel chemical and physical
requirements are found in figure 4-35. The system consists of four
fuel tanks (F-1, F-2, F-3, and F-4), tank pressurization components,
distribution manifolds, control valves, switches, sensors, piping,
interconnect lines, and the connecting hardware required to fill
or drain the tanks, bubble fuel before flight, pressurize the tanks,
and supply fuel to the engines.
The fuel tanks are interconnected at the top by a pressurization
and vent manifold to ensure equal pressurization in all four tanks
and to maintain the required net positive suction head (NPSH)
at the engine fuel pumps. The manifold has two vent valves thatvent the tanks in the event of overpressurization before flight, and
also to vent the tanks during the fuel fill sequence. The system
is not capable of becoming over-pressurized during flight due to
the rapid fuel consumption. The fuel tank sumps are interconnected
by a fuel transfer line assembly and interconnect lines that ensurea uniform fuel level in all four tanks and an equal distribution
of fuel to the engines. In the event of an engine failure, the fuel
normally consumed by the inoperative engine is supplied to the
operating engines. Each tank supplies fuel to one outboard and
one inboard engine through suction lines connected to the tank
sump. A normally open prevalve connected between the tank sump
and the suction line permits control of fuel flow from the tank
sump to the engine. The prevalves provide a backup capability
for fuel shutoff to the main fuel valve in the H-I engines. During
flight, two engine cutoff sensors, one each located in the bottom
of tanks F-2 and F-4, generate a signal when the fuel decreasesto their level that initiates inboard engine shutdown. Similar engine
cutoff sensors in the lox system initiate inboard engine shutdown
if lox depletion occurs prior to fuel depletion. Outboard engine
shutdown occurs approximately 3 sec after inboard engine shut-down. The outboard engines are normally shut down simult-
aneously when engine thrust decay causes the outboard engines
interconnected thrust OK pressure switches to deactuate. However,
if fuel depletion occurs prior to lox depletion, two fuel depletionsensors, one each located in the sump of tanks F-2 and F-4, initiate
outboard engine shutdown when the fuel level reaches the sensors.
Three temperature sensors, located in each fuel tank, monitor the
temperature of the fuel for fuel density calculations and electrically
transmit the results to the facility tanking computer prior to flight.
An overfill sensor, located in the top of tank F-4 sends a signal
to the tanking computer in the PTCS which terminates the fuel
fill sequence in the event of tank overfill. Pressurized GN 2 (290
psig at ground regulator) is bubbled through each fuel suction line
to agitate the fuel and aid in maintaining uniform fuel temperature
in each tank. Fuel bubbling begins just before lox fill and continues
until the start of fuel tank pressurization. The fuel tanks are pres-
RP-1 CHEMICAL AND PHYSICAL REQUIREMENTS
4-30
surized with helium at approximately 2 min and 43 sec before
launch until S-IB stage flight is completed. The tank pressure
ensures the required fuel NPSH at the engine fuel pumps for engine
starting and also to prevent the formation of a vacuum in thetanks as fuel is consumed during flight. Fuel tank pressurization
starts at 29.6 to 32.4 psia and is maintained at this level until
engine ignition. As the fuel is consumed, the tank pressure decreases
to a minimum of 11.5 psia.
Fuel Fill.
The RP-I fuel is stored in the facility storage tanks. Several days
prior to transfer to the S-IB stage, the fuel is processed through
a filter-separator unit that removes water and foreign matter that
may have accumulated. The fuel is then transferred to the S-IB
stage through a cross-country transfer line. Transfer of fuel from
the facility storage tanks include all operations necessary to fill
the S-IB stage fuel tanks. The operations include preparation for
fuel transfer, manual fill, automatic fill, level adjust drain, and
replenish. The manual fill and automatic fill operations are per-
formed approximately 2 days before launch and the level adjust
drain and replenish are performed on launch day.
Preparation for Fuel Transfer. Several major functions performed
before transferring fuel are as follows:
a. The amount of fuel required for this particular mission is pro-
grammed into the PTCS.
b. The fuel filling mast located on the launcher is attached to the
S-IB stage fuel fill and drain nozzle.
c. All closed hand valves in the transfer lines that permit fuel flow
are opened.
d. Electrical power is applied to the RP-1 control panel and
other necessary control components located in the LCC.
e. The pneumatic control pressures are made available at the
facility pneumatic control console located in the fuel storage facility.
f. The S-IB stage fuel vent valves and the fuel fill-and-drain valve
are opened using 750 psig GN 2 control pressure from the pneumaticcontrol console.
Fuel Transfer to the S-IB Stage. Fuel transfer to the S-IB stage
is accomplished as follows:
a. The manual mode of fuel transfer is initiated at the RP-I control
panel.
b. The solenoid valves in the transfer lines are each manually
energized to the open position.
REQUIREMENTS
GRAVITY °API - MINIMUM (MIN)
(SPECIFIC (sp) GRAVITY, MAX)
GRAVITY °API - MAX (sp GRAVITY, MIN)
EXISTENT GUM, MILLIGRAMS (mg) PER
100 MILLILITERS (ml), MAX
POTENTIAL GUM, 16 HOURS AGING rng
PER 100ml, MAX
SULFUR, TOTAL, % WEIGHT, MAX
MERCAPTAN - SULFUR, % WEIGHT, MAX
FREEZING POINT, OF, MAX
THERMAL VALUE: HEAT OF COMBUSTION
BTU/Ib, MIN
VISCOSITY, CENTISTOKES AT-30°F, MAX
AROMATICS, VOLUME %, MAX
OLEFINS, VOLUME %, MAX
SMOKE POINT, MILLIMETERS, MIN
COPPER STRIP CORROSION, ASTM
CLASSIFICATION, MAX
FLASH POINT, MIN
RP-I
42.0 (0.815)45.0 (0.800
7
14
0.05
0.005
-40
18,500
16.5
5.0
1.0
25.0
1
110°F
Figure 4-35
C-H 14394
Section IV S-IB Stage
RP-1 FUEL iYSTEM PRESSURIZATION DIAGRAM
Figure 4-36
oid valves that route the helium from the spheres to the sonic
nozzle. This switch is located in the top of tank F-3, and closes
the solenoid valves when the pressure in the tanks exceeds 32.4
psia. In the event that tank pressure exceeds 35.7 psia, the fuelvent and relief valves mechanically vent the excess pressure to
the atmosphere. At liftoff the electrical circuit to the solenoid valves
and pressure switch is disconnected, and uninterrupted pressure
from the spheres flows through the open solenoid valves and sonicnozzle into the tanks. As the fuel is depleted in S-IB stage flight,
the tank pressure decreases to a minimum of ll.5 psia.
Measurements.
Figure 4-37 lists the fuel system flight measurements and indicates
the information that is displayed.
Fuel System S-IB Stage Components.Detailed descriptions and characteristics of the major components
of the fuel system are presented in the following paragraphs.
CH.1,1636.1
Fuel Fill and Drain Valve. The fuel fill and drain valve .controls
filling and.draining o.f the tanks and is.installed in the fill and
drain line leading to the sump of tank "F-I. The ball-rotor gate
is operated through a rack and pinion gear arrangement and is
driven to the ope n position when pneumatic control pressure is
appiied to the actuator assembly. Simultaneously the rack. and
pinion arrangement compresses the return spring. When pneumatic
control pressure is removed, the ball-rotor gate returns to the
normally closed position. A position indicator switch monitors the
fully open and fully closed positions for remote readout on the
S-IB component test panel in the LCC.
Fuel Prevalve. There are eight prevalves, one each located between
each suction line and the tank sump. The prevalves are normally
open and provide backup capability to the main fuel valves in
the H-I engines for fuel cutoff during flight. The ball-rotor gate
is operated through a rack and pinion gear arrangement and is
driven to the closed position when pneumatic control pressure is
4-31
Section IV S-IB Stage
c. The fuel storage facility transfer pump is energized and fuel
is pumped at approximately 200 gpm into the fuel tanks.
d. When the tanks are 15-percent full, as monitored in the LCC,
the manual operation is terminated. This permits a leakage checkof the transfer lines.
e. The automatic mode of fuel transfer is then initiated at the
RP-1 control panel. The automatic fill functions include fast fill,slow fill, and line drain.
f. The necessary solenoid valves in the facility transfer lines are
automatically energized to the open position.
g. The pump is energized and fuel is routed through the transfer
line at a flow of approximately 2000 gpm into the S-IB stagefill-and-drain line and into the fuel tanks.
h. The tanks are filled to 98-percent full and then the tanking
computer automatically initiates the slow fill sequence. The slow
fill sequence restricts the fuel flow to 200 gpm. The percent offill is monitored in the LCC.
i. The tanks are slow filled to approximately 102 percent, then
the fill sequence is automatically terminated and the transfer line
drain operations are automatically sequenced.
j. Fuel fill and drain valve control pressure is removed and the
valve returns to the normally closed position.
k. Fuel in the facility transfer lines is returned to the RP-I storagetanks.
I. Immediately after the line is drained, a 750-psig GN 2 transfer
line and fuel fill mast purge is automatically sequenced.
m. After lox has been loaded and just prior to launch, a fuel level
adjust drain operation is performed. When lox is in the S-IB loxtanks, heat transfer between the lox and fuel tank occurs and cools
the fuel. The fuel density increases and the fuel load has to be
adjusted to the required mass for a given mission.
n. The corrected fuel load is programmed into the facility tanking
computer and the level adjust drain operations are initiated at the
RP-I components control panel.
o. The fuel fill and drain valve on the S-IB stage is opened and
the excess fuel is drained into the facility transfer line until the
fuel level in the S-IB stage tanks decreases to the 100 percent full
level (see paragraph q). The fuel fill and drain valve returns to
the normally closed position.
p. The facility transfer line is drained automatically and the GN 2
purge of the line and fuel fill mast is again performed.
q. In the event that the fuel tanks are less than 100 percent full,based on the latest fuel mass requirements, a replenish operation
must be performed.
r. The fuel fill and drain valve and the fuel vent valves on the
S-IB stage are again opened and fuel at 1000 gpm is routed through
a fast fill valve and starts filling the transfer line. A 30-sec time
delay is initiated and closes the fast fill valve when the transferlines are full. The fuel is then routed through the slow fill line
at 200 gpm into the S-IB gtage tanks.
s. When the tanks are again 100-percent full, the replenish opera-
tion is terminated and the line drain and purge functions are
automatically performed.
Fuel Drain From S-IB Stage To Facility Storage.
In the event of launch cancellation and the fuel has to be drained
from the S-IB stage, a drain operation is performed as follows:
a. Drain operations are initiated at the RP-I components control
panel and drain functions are automatically sequenced.
4-32
b. The S-IB stage fuel tanks are pressurized to flight pressure from
the facility helium source and valve panel no. 9 through the S-IB
stage fuel tank pressurization components.
c. The S-IB stage fuel fill-and-drain valve and the necessary facility
transfer line valves are opened and the fuel drains to the facility
storage tanks.
d. When the fuel level decreases to 10 percent of the total fuel
load, the tanking computer initiates a signal that sequences the
line drain functions to return fuel in the facility transfer lines to
the RP-1 storage tanks.
e. Immediately after the line is drained, a 750-psig GN 2 transfer
line and fuel fill mast purge is automatically sequenced.
f. The fuel fill and drain valve is returned to the normally closed
position.
Fuel Bubbling.
Pressurized GN_ at approximately 135 psig is routed from valve
panel no. 10 in the ML to the S-IB stage and bubbles the fuel
in the fuel suction lines and the tanks. GNz flow into the suction
lines and subsequently through the fuel tanks tends to maintain
a uniform fuel temperature within each tank. Fuel bubbling is
first initiated from the S-IB firing preparation panel in the LCC
approximately 8 hr and 45 min before launch and is stopped during
level adjust drain operation and then started approximately 10
min prior to launch and continues until fuel tank pressurization
begins. The GN z is routed through a ring line manifold and branch
lines to each fuel suction line. Each branch line is provided withan orifice to control GN 2 flow and a check valve to prevent reverse
fuel flow into the GN 2 bubbling line. The bubbled GN 2 rises
through the suction lines and the normally open prevalves to the
ullage area of the fuel tanks and is vented through the open fuelvent valves.
Fuel Tank Pressurization.
The fuel tanks are pressurized (figure 36) with helium at 29.6 to
32.4 psia starting from approximately 2 min and 43 sec prior to
launch and continuing until S-IB flight is completed. The tank
pressure maintains a pressure head (NPSH) for starting the engine
fuel pumps and provides structural support by preventing theformation of a vacuum in the tanks as fuel is depleted during
flight. Pressurizing components of the S-IB stage include two high
pressure storage spheres, solenoid valves, pressure switches, a sonic
nozzle, and distribution lines. Prior to fuel loading, the two high
pressure storage spheres are prepressurized to approximately 1600
psig. The helium is supplied from valve panel no. 9 through a
Q-D coupling (service arm no. IA) connected to the upper umbilical
on the S-IB stage. The helium enters the stage and passes through
a filter and check valve into the two storage spheres. During the
tank pressurization sequence, the storage spheres are fully pres-
surized to 3000 psig and are continuously replenished if necessary
prior to launch. From the spheres the helium is routed through
two normally open solenoid valves and a sonic nozzle into thedistribution lines to each tank. The sonic nozzle meters the helium
flow and maintains a constant rate of flow into the tanks. Prior
to launch, a pressure OK switch located on the outlet of one sphere
monitors the sphere pressure and actuates and deactuates to cause
a solenoid control valve in the facility source to shut off or open
the supply. The switch actuates at 2965+30 psia on increasing
pressure to shut off the supply and deactuates when the sphere
pressure decreases to 2835 psia. The FUEL PRESSURIZATION
PRESSURE OK readout is in the LCC. Another pressure switch
monitors tank pressure and controls the operation of the two solen-
FUEL YSTEMS MEASUREMENTSNUMBER
XC179-F1
XCI79-F2
XC179-F3
XC179-F4
XD2-F3
VK17-F2
i VK 18--F4
L20-F!
L20-F3
NOMENCLATURE
TEMPERATURE, FUEL
TEMPERATURE, FUEL
TEMPERATURE, FUEL
TEMPERATURE, FUEL
PRESSURE, HE, FUEL TANK
LEVEL CUTOFF, FUEL
LEVEL CUTOFF, FUEL
LEVEL, FUEL, DISCRETE
LEVEL, FUEL, DISCRETE
AUXILIARY DISPLAY
AUXILIARy AND MCC-H DISPLAY
ESE DISI_.AY
RANGE
+,to°c
-_O +40°C
+o+4o°c
+o+_°c
+0 +45 PSIA
ON OFF
ON OFF
ON OFF
ON OFF
DISPLAY
Figure 4-37
C-H 14542
applied to the actuator assembly. Simultaneously the rack and
pinion arrangement compresses the return spring. When pneumatic
control pressure is removed the ball-rotor gate returns to the nor-
mally open position. A position indicator switch monitors the fully
open and fully closed positions for remote readout on the S-IB
firing preparation panel in the LCC. All prevalves in the OPEN
position complete an interlock for the start of automatic launch
sequence.
Engine Cutoff Fuel Sensor. Two engine cutoff fuel sensors initiate
an electric signal to shut down the inboard engines when the fuellevel in tanks F-2 or F-4 falls below the level of the sensors.
Fuel Depletion Sensors. Two fuel depletion sensors initiate an elec-
trical signal to shut down the outboard engines when the fuel in
tanks F-2 or F-4 is depleted below the level of the sensors.
Fuel Overfill Sensor. The fuel overfill sensor initiates an electrical
signal to the PTCS to terminate the fuel fill sequence should thefuel level in the tanks rise to the level of the sensor. A FUEL
OVERFILL readout is monitored on the S-IB component test panelin the LCC.
Fuel Temperature Sensor. Three temperature sensors in each of
the four fuel tanks are located parallel to the longitudinal axis
of the tanks. The temperature sensors monitor the temperature
of the fuel and electrically transmit the results to the facility fuel
tanking computer. The temperature data is used in density calcula-
tions for programming the correct fuel mass to be loaded into the
S-IB stage fuel tanks.
High Pressure Storage Sphere. Two high pressure storage spheresare used to store the helium, 19.28 ft _ each, required for inflight
fuel tank pressurization. The storage spheres are pressurized to
3000 psig from the facility source prior to launch and are maintained
at this pressure until liftoff. After liftoff these spheres pressurize
the fuel tanks during S-IB stage flight. Each of the two identical
spheres is formed by welding together two fully-machined forged
titanium hemispheres. The spheres are proof tested at 4650 psig
through a temperature range from -125 ° F to 200 ° F and are
designed to withstand 6200 psig without rupture.
Solenoid Operated Control Valve. Two normally open solenoid
operated control valves control the fuel tank pressurization prior
to liftoff. A pressure switch located on top of tank F-3 senses the
correct tank pressure and causes the solenoid valve to close or
open to maintain tank pressure of 29.6 to 32.4 psia. At liftoff theswitch is disconnected and the solenoid valves return to the normally
open position. Helium then flows at the maximum flowrate through
Section IV S-IB Stage
the sonic nozzle into the fuel tanks. During sphere fill or at any
time prior to launch, the valves can be closed by operation ofthe FUEL PRES'G VALVES NO. 1 and NO. 2 switch on the
S-IB component test panel in the LCC. During venting, prior tolaunch, an electrical signal from the facility closes the valves until
tank pressurization is required again.
Pressure Switch. The pressure switch controls the operation of the
solenoid operated control valves to maintain the fuel tank pressure
at 32.4 lasia (max) prior to liftoff. When the fuel tank pressure
is below 29.6 psia (min) the switch causes the fuel pressurizingcommand circuit in the LCC to energize and the solenoid operated
valves return to the normally open position and the helium is routed
to the tanks. When the tank pressure increases to 32.4 psia (max),
the switch causes the fuel tanks pressurized circuit in the ESE to
energize and the solenoid operated valves close and the helium
flow to the tanks is stopped. The switch also causes the FUEL
PRESSURIZED indicator to light on the S-IB firing panel when
the tank pressure is 29.6 to 32.4 psia. The switch has no inflightfunction.
Vent Valves. Two vent valves vent the fuel tanks during fill opera-
tions or when emergency venting is required. The normally closed,
spring-loaded, poppet-type valves are opened by pilot valves. GNz.
from the facility source opens the valves during fuel filling opera-tion. Pressure sensing lines from tanks F-3 and F-4 route ullage
pressure to provide operating pressure for emergency venting. The
valves are provided with pos_ion indicators for remote readout
on the S-IB component test panel and the S-IB firing panel inthe LCC.
Fuel Pressurization Supply OK Switch. The fuel pressurization
supply OK switch monitors the helium stored in the two high
pressure storage spheres. The switch is set to actuate at 2965-+30
psia increasing pressure and to deactuate at 2835 psia minimum
as pressure decays. When the spheres are fully pressurized, the
switch actuates and the power is removed from the facility control
valve which shuts off the helium supply. Also, a FUEL PRES-
SURIZING PRESSURE OK readout signal is transmitted to the
S-IB firing preparation panel in the LCC. Should the pressure decayand cause the switch to deactuate, the functions reverse to remove
the pressure OK readout, apply power to the facility control valve,and recharge the storage spheres. The switch has no flight function.
CONTROL PRESSURE SYSTEM.
The S-IB stage control pressure system located in the aft skirt
of fuel container F-3 stores 3100-psig GN 2 and supplies 750-psig
regulated pressure for operation of fuel and lox prevalves, calorim-
eter purge, gearbox pressurization, and lox pump seal purge. Pres-
sure switches and transducers monitor system conditions and pro-
vide inputs to the S-IB stage telemetry system. GN z from valve
panel no. 10 enters the control pressure system through short cable
mast 4 (figure 4-38) and pressurizes the l-ft _sphere. During preflight
test activities, the control pressure system receives 50-psig GN 2
to purge the system. Adjustment of the regulator for outlet pressure
of 40 to 50 psig and opening the gearbox pressurization and lox
seal purge hand valve permits GNz flow through the systems. This
operation continues for a minimum of 30 min before pressurizing
the sphere to prelaunch operation pressure of 1500 psig. Approxi-
mately 5 hr before launch the sphere pressure is increased to 3100
psig. An orifice in the supply line in cable mast 4 valve panel
controls the sphere pressurization rate. After the sphere pressureincreases to 3100 psig, the 30001b control pressure OK signal from
the high pressure OK switch energizes a solenoid valve that bypasses
the orifice to permit sufficient GNz flow to replenish the sphere.
Replenish continues until the time for ignition signal at T-3 sec.
The high pressure OK switch actuates at 2965---+30 psia and provides
4-33
Section IV S-IB Stage
;-IB STAGECONTROLPRESSURESYSTEMDIAGRAM
TO LOX VENT
AND RELIEF
VALVE
3-WAY
VALVE
OK SWITCH
DISCONNECT
• COUPLING
CHECK FILTER
VALVE
SOLENOID 9r "--.=0 + 50 PSIG VENT
CONTROL VALVE It RIRELIEF VALVE VALVE AND
MANIFOLD
t ASSV--XCF _ ONTROL PRESS FILTER
PRESSURIZATI
,_ ANDLOX_ I I ISEAL ROE III
GROUND PREVALVE
CONTROL LINE -_
1--11--I
ORIFICE
(TYP) -_
QUICK
DISCONNECT
METER_ TO FUEL TO LOXPREVALVE PREVALVE
(TYP 8 (TYP 8
CALORIMETER PLATES) PLACES)
CALORICH-14628-1
Figure 4-38
an output to illuminate the CONTROL 3000 LB OK indicator
on the S-IB firing prep panel. After liftoff the switch has no function.
Pressure transducer XD40-9 provides continuous monitoring of
sphere pressure before liftoff and during flight. Measurement prefix
X indicates that the signal is routed to KSC/LVO-1 for real-time
analog recording. See figure 4-39. Redline value for XD40-9 is
3300 psia maximum.
Sphere pressure is regulated to 750-4-15 psig and distributed to
the vehicle systems through the control pressure manifold. The
control pressure OK switch actuates at 625-----25 psig and provides
an output that illuminates the CONTROL 750 LB OK indicatoron S-IB firing prep panel. A 950±50 psig relief valve provides
overpressure protection for the 750 psig system. Pressure transducers
XD41-9 and XD42-9 provide continuous monitoring of the regula-
tor output pressure before liftoff and during flight. XD41-9 and
XD42-9 redline values are 710 psia minimum and 815 psia maxi-
mum. This information requires periodic-monitoring from 3100-
psig pressurization to the control pressure sphere until automatic
sequence start.
TEST DATA.
The l-ft _ fiberglass storage sphere is proofed at 5000 psig and has
a minimum burst pressure of 6660 psig. The operating pressure
is 3000 psig with an operating temperature of -65 to + 160 ° F.The most critical items in the system are the bottle fill-and-vent
valve, the 750-psig regulator, and the relief valve. Failure of the
fill-and-vent valve to remain closed would result in a probable
mission loss. Failure of the regulator or relief valve would result
in a possible mission loss. The control pressure system components
have been qualified and have flown previous Saturn flights. Prior
to qualification testing, each component of the qualification test
sample was individually proof tested. The sphere was pressurized
hydrostatically; all other components were pressurized with helium.
Each component was pressurized to 150 percent of normal operating
pressure and maintained in that condition for 5 min. During the
final 2 min of the 5 min period a leakage check was performed.
The pressure was reduced to zero and an examination of the
components for distortion was made. No discrepancies were de-tected.
CONTROL PRESSURE SYSTEM FUNCTIONS.
fhe calorimeter purge control solenoid valve may be energized
by the CALORIMETER PURGE switch on the S-IB firing prep
panel or, during countdown by the power transfer command. GN z
flows through two orifices that reduce the flowrate to 0.218±0.022
lbm/min/calorimeter. Ambient purge prevents accumulation of
combustion products and other contaminants on the calorimeter
radiation window; purge continues throughout S-IB stage operation.The calorimeters measure heat flux in the engine area and relay
the information through TM. The measurements are of prime
interest from liftoff until S-IB stage impact. See figure 4-39.
Gearbox pressurization and lox pump seal purge begins when the
control system receives initial pressure. A three-way valve permits
the purge to be shut off when necessary during checkout. An orifice
reduces the GN 2 flow to eac h engine to 0.23±0.09 lbm/min. The
4-34
purge line tees and one branch becomes the lox pump seal purgewith a flowrate of 1.35 scfm and the other branch becomes the
gearbox pressurization with a flowrate of 1.85 scfm. A check valve
in the gearbox pressurization line prevents reverse flow into the
purge line. Gearbox pressure of 2 to 10 psig is maintained during
flight. See H-I Engine, for additional information.
The 750-psig control pressure closes the prevalves at H-1 engine
cutoff, serving as a backup propellant flow termination device with
the main propellant valves on the engine. A normally closed solen-
oid valve for each pair of normally open fuel and lox prevalves
prevent closure until deactuation of TOPS on each engine. The
solenoid control valve then opens and pressurizes the closing actua-
tors of fuel and lox prevalves. See Lox System and Fuel System
for additional prevalve information.
During flight, sphere pressure will decay because of continuous
purging of the calorimeters, lox pump seals, and pressurizationof the turbopump gear boxes. See figure 4-40 for predicted decay
rates based on available pressure at liftoff.
l[ll'lll i '_ 11.1,111111,11.111i III
MEAS
NO TITLE
C603-6 FLAME SHIELD TEMP (R. CAl.)
C609-3 HEAT SHIELD TEMP (R. CAL)
XD40+9 CONTROL EQUIP SUPPLY PRESS
XD41-9 CONTROL EQUIP REGULATED
PRESS
XD42-9 CONTROL EQUIP REGULATED
PRESS
0 TO 115 W/cm 2
0 TO 45 W/era 2
0 TO 3500 psia
0 TO 1000 psio
0 TO 1000 psia
Figure 4-39
CH.14392-1
Section IV S-IB Stage
SYSTEM FLIGHT HISTORY.The S-IB-I control pressure system sphere pressure at ignition was
2920 psi. Minimum acceptable pressure, determined by the deac-
tuation setting of the high pressure OK switch, was 2862 psi. The
regulated pressure ranged between 760 and 768 psi well within
the specified redline value of 710 to 815 psi. The predicted sphere
pressure decay compared favorably with the actual sphere pressure
decay. Actual sphere pressure at 2 min 30 sec flight time was within
25 psi or 2.5 percent of predicted value. The S-IB-2 control pressure
system utilized two 1-ft _ spheres because of increased pneumatic
requirements for supplying purge to a fifth calorimeter. Sphere
pressure at T-3 sec was 3160 psia. The regulated pressure was wellwithin limits throughout countdown and flight. At T-10 sec, the
regulated pressure was 768 psia and then varied between 760 and
770 psia during flight. The S-IB-3 sphere pressure at T-10 sec was
3040 psi. Regulated pressure ranged between 770 and 775 psi during
flight. Two 1-ft 3 spheres were used in the S-IB-3 control system
to accommodate the additional purge flowrates for five calorimeters
and one spectrometer. From T-10 sec until outboard engine cutoff,
the pressure declined steadily to 1820 psi. Final sphere pressure
was within 50 psi of predictedvalue. At approximately 2 min 20
• sec into flight, a change in slope of the decay curves was caused
by the pneumatic requirement for dosing the prevalves. The regu-
lated pressure was 770 psi by 2 min 30 sec into flight and remainedwithin the range of 710 to 815 psi. S-IB-4 and S-IB-5 each utilized
a single 1 ft 3 sphere for control pressure storage. On each of these
flights, the regulated pressure and gas usage rate remained within
acceptable limits. The S-IB-4 regulated pressure varied between
770 and 785 psi; S-IB-5 varied between 759 and 754 psi.
I
ua
3200
3OOO
2800
26OO
24OO
2200
2ooo
1800
1600
1400
1200
* NOTE: CONTROL PRESSURE DECAY BEGINS
WHEN REPLENISH TERMINATES AT
T-3 SEC (IGNITION COMMAND).
MINIMUM ACCEPTABLE PRESSURE AT
T-3 SEC: 2835 )s|.
*pOSSIBLE PATHS OF
PRESSURE DECAY
DURING FLIGHT
lOOO0:30 1:00
RANGE TIME (MIN:SEC)
Figure 4-40
1:30 2:00 2:30
C.H 14393.1
4-35
SectionIVS-IBStage
HYDRAULI( SYSTEM
FLUID;FILLSYSTEMpRECHARGEWITHACCUMULATOR60C22311HYDRAULICwITH FLUID ...... _':::×:':" J HYDRAULICO.2TO.0OPSG LOWPRESSUBE-- SERVArORSO PCO_OLVOLTAOESOKrH_ FLU,° J-,,oKTOSTA_HYD_UUC_MPS" ELECTRICAL
L,GHTW,LLGLOW. _ FL,GHTCONTROL\COMPUTERIINlU) //L-.-_HI I
('_POSmONINGTHEHYDRAULICPUMPSENABLE ENGINE \ _ // (L_it I I"_sWnCH (ss)"ON" WILLENABLETHEAUXtUARY -"'--"_\ /-- TURBOPUMP _. lET! _ _ _!i] I I
PUMPS CONTROL CIRCUIT AND LIGHT A PANEL _ -- _,_' A_v _IV_ _ _ - _-----,,..=_,_1_/__ j jINDtCATOR "_'_ .... __. % - _ I I
• c o_
C_POSITIONING THE SYSTEM SWITCHES ($1 F _4"4A/D_THROUGH S4)"ON" SUPPLIES AC POWER TO / ./_ _ J l
THEAUX,L,ARYPUMPMOTO,S.TH,S,SDONE.> _ _ _O_"_[__4 MIN 30 SEC BEFORE LAUNCH.
(_')THE AUXILIARY PUMP SUPPLIES HIGH PRESSURE PUMP_FLUID TO THE ACCUMULATOR-RESERVOIR AND
MANIFOLD ASSY.
WtTH THE SYSTEM THUS PRESSURIZED, THESERVOACTUATORS HOLD THE ENGINE IN THE
DESIRED POSITION.
QWHEN THE ENGINES START, THEY DRIVE THE
MAiN PUMPS PROVIDING HIGH PRESSURE
FLUID TO THE ACCUMULATOR-RESERVOIR AND
MANIFOLD ASSY.
THE "ALL ENGINES RUNNING" SIGNAL DIS-ABLES THE AUXILIARY PUMPS CONTROL CIR-
CUIT REMOVING ELECTRICAL POWER FROM
ALL FOUR AUXILIARY PUMP MOTORS.
THE FLIGHT CONTROL COMPUTER PROVIDESATTITUDE CORRECTION COMMANDS TO THE
SERVOACTUATOR SERVOVALVE.
G THE SERVOVALVE CONTROLS FLUID PRESSURETO RETRACT OR EXTEND THE SERVOACTUATOR
PISTON THEREBY GIMBALING THE ENGINE.
A POTENTIOMETER ON EACH SERVOACTU-ATOR PROVIDES POSITION FEEDBACK TO THE
FLIGHT CONTROL COMPUTER.
BLEED
VALVE
PLACES)-
AUXILIARY
OHIGH PRESS.
Q-D (GROUND FILL) --/
DIFFERENTIAL
PRESSURE _,,INDICATOR _
[_ DISR.AYED THROUGHODAS ON DISPLAY RE-
CORDERS IN THE LCC
[_ DISPLAYED THROUGHDDAS ON S-IB HYDRAULIC
PANEL AND DISPLAY '
RECORDER IN THE LCC
AND AT MCC-H
SERVOVALVE
TRANSDUCER ((_'_
VXO29-1, 0 TO _ \
4000 PSIG _
PRESS. Q-D
(GROUND RETURN)
j ,HYO.AULCPUM ENABL ONII_ 400 Hz 3 PHASE
200V (LINE TO LINE)4e_.
ONm ON
I s-zaHYDRAULICI
L-
AUTO
S4 OFF S5
CH.14_7-2
4-36Figure 4-41
S-IB HYDRAULIC SYSTEM.
The flight control computer in the IU combines attitude error
signals from the navigation and guidance systems with rate signalsfrom control accelerometers and rate gyros and produces attitudes
correction commands. Servoactuators on the outboard engines re-
spond to these commands by gimballing the engines, thereby di-
verting the thrust vectors to correct control errors or execute pro-
grammed pitch and roll maneuvers. The gimbal direction and rate
are proportional to the polarity magnitude of the command signal.
Each of the four outboard engines is equipped with an independent
closed-loop hydraulic system. Each system consists of five majormodules: accumulator, reservoir, and manifold assembly; main
• pump; auxiliary pump; pitch servoactuator; and yaw servoactuator.
Tubing and flexible hoses interconnect these modules. Components
within modules are ported through manifolds. This keeps external
plumbing at a minimum reducing leakage problems and making
the system less vulnerable. When the servoactuators extend and
retract, the engine can be gimbaUed a maximum of ±8 deg from
the null position in both the pitch and yaw axes. In the null position,
each engine is canted 6-deg outward from the vehicle centerline.
FILLING.
Normally the hydraulic system should not require filling and bleed-
ing after the booster leaves the Michoud assembly plant. However,
Section IV S-IB Stage
if the system is opened, filling and bleeding must be repeated as
follows. Prior to filling the system with hydraulic fluid, the ac-
cumulator is charged to 1600 psig with GN 2 from the launcher
pneumatic manifold through the GN 2 precharging valve. Eachaccumulator takes 0.251 Ibm of nitrogen. Hydraulic fluid is then
pumped into the system from the hydraulic servicer cart through
the high pressure, quick-disconnect coupling. The system is thenbled to the reservoir desired level as monitored by the fluid level
potentiometer. Excess fluid returns to the servicer cart through the
low pressure quick-disconnect coupling. After filling, the GN 2
precharge is bled down to 10 psig if the system is to remain
inoperative for 10 days or more.
PRELAUNCH PREPARATION.
Just before operating the system for checkout or during countdown,the accumulator is charged with GN 2 to 1600 psig. If the pump
motor temperature is under 350+---18 ° F and the low pressure fluid
temperature is under 200±10 ° F on all four systems the MOTOR
TEMP OK and FLUID TEMP OK lights on the S-IB hydraulic
panel will illuminate (figures 4-41 and 4-42). If the control voltage
is OK, the OK TO START HYD PUMPS light will illuminate.
To start the systems operating for checkout or countdown, the
enable switch is placed on the ON position and the HYD PUMPS
ENABLE ON light illuminates. Placing each system switch in the
ON position supplies facility ac power to each auxiliary pumpmotor. The auxiliary pump can deliver 2.2 gpm at a minimum
HYDRAULIC SYSTEMDIAGRAMMECHANICAL DRIVE
FOR H-I ENGINE
SYSTEM BLEEDVALVE VALVE
MAIN
SEEPAGE
PLUG
L,P.
RELIEF
L.P.
ISOLATION BLEED
CHECK VALVE VALVE
IACTUATOR (2)
H.P.
RELIEF
TRANSDUCER
IXILIARY
PUMP
ISOLATION
CHECK VALVE
SEEPAGE
AUXILIARY
PUMP
H.P.BLEED
DIFFERENTIAL
PRESSURE
INDICATOR
SWITCH
CHARGING
VALVE
HIGH PRESSURE
LOW PRESSURE
Figure 4-42
VALVE
CH 14627
4-37
Section IV S-IB Stage
RESERVOIR
BLEEDER
RESERVOIR
PISTON
AUXILIARY PUMP i
L.P. RETURN PORT---_
BOOTSTRAP \_//_
SLEEVE _
ACCUMULATOR CAP J_
_GN 2 CHARGING
VALVE
BLEEDER
VALVE
GUIDE
NUT
ACCUMULATOR-RESERVOIR AND MANIFOLD CHARACTERISTIC ¢
LOCATION
TEMPERATURE
OPERATING
PRESSURE
ACCUMULATOR
OPERATING
PROOF
BURST
GN 2 PRECHARGE
RESERVOIR
OPERATING
PROOf
BURST
FLUID
ACCUMULATOR
ACTIVE
TOTAL
RESERVOIR
ACTIVE
TOTAL
STATION 67, ON EACH
OUTBOARD ENGINE
-4Oto +275OF
3200 psig NOMINAL
5000 psig MIN
7500 psig MIN
1600 pslg
53.3 pslg
300 psig
600 pslg
60C22311
32 in. 3
38 in.3
104 in.3
122 in. 3
AUXILIARY PUMP
DRAIN INLET_
F,LTER __
I ELEMENT
SYSTEM irjl ii _
• _ACCUMULATOR /
PISTON /
r,°5,!,7::j
TRANSDUCER_
OTENTIOMETER ACTUATOR
_
LOWPRESSURE/ ..._QUICK DISCONNECT--"
HIGH PRESSURE
LOW PRESSURE
FILTER
VALVE
PUMP H.P. IN
VALVE
MAIN
PUMP H.P. IN
PLUG
IIIII
"illJ MANIFOLD
, / R%uIOR
- _ OUTLET
J TO MAIN
PUMP
LOWER
/ MAN_POLD
Note
ACCUMULATOR SECTION OF
MANIFOLD WITH HIGH PRESSURE
RELIEF VALVE IS NOT SHOWN
Figure 4-434-38
pressure of 2850 psig to allow system checkout and to hold the
engines in the neutral position at ignition. As each engine ignites,
its main hydraulic pump, driven by the engine turbopump, begins
delivering high pressure fluid. After the ignition sequence, the all
engines running signal removes power from the auxiliary pumpmotors.
OPERATION.
The main operating pump draws fluid from the low pressure,
53-psig, side of the accumulator-reservoir and increases the pressure
to a maximum of 3250 psig. Inlet pressure is dependent on pump
outlet pressure (60:1). When the auxiliary pump is operating, the
pressures are 50 and 3050 psig (max). These maximums are no-flow
pressures. In the complete system, operating pressure is less than
this value. The fluid then passes through high pressure tube as-
semblies to the manifold assembly and into the high pressure
portion of the accumulator-reservoir. From the opposite manifold
through tube assemblies and flexible hoses, the fluid passes to the
servoactuators. Electrical commands from the flight control com-
puter in the IU signal the servovalves to throttle the fluid in a
direction and at a rate corresponding to the polarity and magnitude
of the command. The high pressure fluid flows to the proper side
of the servoactuator piston causing it to extend or retract. Low
pressure fluid returns to the low pressure side of the accumulator-
reservoir through tube assemblies and flexible hoses and then
through tube assemblies to the pump inlet.
Section IV S-IB Stage
ACCUMULATOR-RESERVOIR.
The accumulator portion, figure 4-43, stores some high pressure
fluid during system operation to help meet sudden servoactuator
demands and dampen pump surges. A floating piston allows the
accumulator volume tovary according to the quantity of fluid stored
at any instant by means of GN z precharge on the piston side
opposite the high pressure fluid. The reservoir portion stores low
pressure fluid to feed the pump inlets. A bootstrap piston in the
reservoir has one piston surface acting on the low pressure fluid
and another piston surface, one-sixtieth the area of the first, exposed
to the high pressure fluid in the accumulator. The high pressure
fluid forces the piston down on the low pressure fluid, maintaining
reservoir pressure at 1/60th of accumulator pressure (47-53 psig)
to help prevent pump cavitation. The volume of fluid in this unit
and the pump induced pressure is sufficient to meet the maximum
expected servoactuator demands with a safety factor of 1:6.
MANIFOLD.
One portion of the manifold, figure 4-43, contains the main system
filter, which cleans the fluid as it flows from the pumps to the
accumulator. Another filter cleans the auxiliary pump case drain
fluid. Check valves protect against reverse flow in the outlet tubes
of each pump. The differential pressure indicator shows, during
prelaunch operations, when the main filter needs cleaning. Another
portion of the manifold contains high and low pressure relief valves
MAIN PUMP
RELIEF
PORT
CYLINDER
INLET
ROLLER
BEARING PLATE
OUTLET
HOLD DOWN
PLATE
Figure 4-44
ADJUSTMENT
SCREW
VALVE
RATE
'PISTON
HANGER
PISTON
MAIN PUMP CHARACTERISTICS
STATION 80LOCATION
TEMPERATURE
OPERATING
PRESSURE
ZERO FLOW
FULL FLOW
INLET
SPEED
RATED FLOW
LUBRICATION
-65to +275°F
3200 + 50 psig
2900 psTg
7 to 50 psig
4300 rpm
17 gpm @ 4300 rpm
SELF-LUBRICATI NG
GH 14246. I
4-39
Section IV S-IB Stage
to protect both portions of the system from overpressurization. The
manifold contains the high and low pressure quick-disconnect
couplings used to fill the system.
MAIN PUMP.
This two-stage, cam-actuated, variable displacement pump (figure
4-44) uses a sensing of output pressure to compensate for fluctua-
tions, thereby maintaining a constant output pressure of 3200 psig.
The pump is in effective operation by the time its cylinder barrel
completes one revolution. Rotation of the cylinder barrel moves
the seven dual-diameter pistons along the variable angle cam plate
(wobble plate) causing the pistons to reciprocate within the cylinder
block. The intake strokes of the pistons cause hydraulic fluid to
be drawn through the inlet port into a fixed pintle and then into
the primary stage of the pump. The discharge strokes of the largediameter portion of the piston pressurize the fluid to approximately
100 psig and route it along the outside of the pintle to the second
stage. (The flow from the first stage is more than is required by
the second stage. Excess fluid is diverted through a relief valveand is routed back to the pump inlet.) In the second stage, the
smaller diameter portion of the pistons pressurizes the fluid to the
operational level of 3200 psig. The high pressure fluid is discharged
as the cylinder block rotates and aligns each piston with the high
pressure outlet port.
Pump outlet can be varied by changing the angle of the cam plate.This is controlled by the rate piston and the stroking piston. Under
high load conditions, the cam plate is held immobile in the maxi-
mum displacement position by the rate piston and pumping load.
During operation these forces are opposed by the stroking piston,which receives control pressure fluid from the compensator valve.
This normally closed valve is sensitive to pump discharge pressure
and is screw-adjusted to maintain outlet pressure at the desiredlevel.
During pump start, when rising discharge pressure approaches 3200
psig, the sensed force on the compensator valve spool overrides
the preset valve adjustment and the spool is displaced, delivering
high pressure fluid to the stroking.piston. The stroking piston causes
the trunnion hanger to rotate, and the cam plate moves to a reduced
angle. As operating conditions stabilize, the cam plate will assume
a relatively permanent position, subject only to minor changes when
output pressures vary because of system demands. A hollow epoxy
seepage plug, filled with sponge, absorbs any leakage around the
drive shaft seals. The seepage plug threads into the base of the
pump. A case drain line routes fluid that has escaped past the
pistons and cylinder barrel back to the pump inlet. The main pump
obtains its power from the engine turbopump gearcase.
AUXILIARY PUMP.
The auxiliary pump (figure 4-45) is a fixed-angle, nine-cylinder,
pressure-compensated unit with variable delivery, controlled by arotor, automatically rotated to meet system requirements of high
pressure hydraulic fluid. The pressure capability is 3000---+50 psig
at zero flow to full flow with 2850 psig minimum. The flowrate
DOUBLE
UNIVERSAL
ROTC
4-40
AND PISTON
ROD
DRIVER ASSY
BLocK
SPACER
VALVE BLOCK
Figure 4-45
AUXILIARY PUMP CHARACTERISTICS
STATION 73LOCATION
TEMPERATURE
OPERATING
PRESSURE
ZERO FLOW
FULL FLOW
INLET FLOW
PROOF
SPEED
RATED FLOW
LUBRICATION
0 to +275°F
300O psig
2850 psig (mln)
37 psia @ 10,000 mm
4500 psig
10,500 rpm NOMINAL
3.0 gpm@ 11,000 rpm
SELF-LUBRICATING
GH 14245-1
m
AUXILIARY PUMP MOTOR CHARACTERISTICS
LOCATION
TEMPERATURE
OPERATING
ELECTRICAL
VOLTAGE
SPEED
POWER RATING
SWITCH
INDICATION
STATION67
0 to 165°F
200 V LINE-TO-LINE,
400 Hz, 3 PHASE
11,300_+ 5% ppm at
25 in-lb TORQUE
3356 W (4.5 HP)
MAX CONTINUOUS
DUTY
OPEN ABOVE 350°F
CLOSE BELOW 310°F
C-H 14244-1
Figure 4-46
at 11,000 rpm is 3.0 gpm minimum. During checkout the flowrate
is limited to 2.2 gpm for 2 min duration by controlling the gimball-
ing rate, thus preventing overheating the auxiliary pump motor.
The pump is operated during the hydraulic system fill and purge
operations, during leak checks, during engine gimballing precheck
operations, and during launch countdown to prepressurize the
system before engine ignition. The pump serves no flight function.
The auxiliary pump is powered by an electric motor (figure 4-46)fed from facility source of 200-V, 3-phase, 400 Hz. The motor is
a 3-plaase "Y" wound unit, capable of developing 11,300 rpm -+'5percent at 4.5-hp load. The pump and motor are mounted on
HYDRAULIC SERVOACTUATOR CHARACTERISTICS
LOCATION
EMPERATURE
_RESSURE
OPERATING
PROOF
BURST
EFFECTIVE PISTON AREA
LOAD
RATED
STALL
STROKE
TOTAL
LIMITED
.ENGTH
ADJUSTMENT
VELOCITY(Input 12 MA dlff)
LOADED (10,000 Ib)
NO LOAD
LUBRICATION
SERVOVALVE
TORQUE MOTOR
2 COILS
COIL RESISTANCE
DtFF INPUT SIGNAL
(RATED)
COIL CURRENT
(MAX)
STATION 76
-65 TO +275 °F
3OO0 prig
4500 pslg
7500 ps[g
5 in 2
10,000 Ibf
15,000 Ibf
9.560 + 0.06 in.
7.640 + 0.06 in.
40 Tn.
0.5 in.
8.75 + 0, -1.75 [n/sec
15.0+ 0, -3.0 in/see
SELF-LUBRICATING
POLARITY OF SIGNAL
DETERMINES COIL USED
1000 _+50 OHMS
12 ± 0.12MA
20 MA
TORQUE
MOTOR AND
FLAPPER
SERVO
VALVE ---_
RETURN
PORT_
PREFILTRATION
VALVEu_PRESS
PORT_
Section IV S-IB Stage
opposite sides of a bracket located on the engine. The motor has
an integral thermal switch that is interlocked in the motor start
circuit and completes the circuit to the MOTOR TEMP OK indica-
tor light on the hydraulic panel in the LCC.
SERVOACTUATOR ASSEMBLY.
Gimbal commands are received from the flight control computer
to actuate the electrohydraulic servovalve in each actuator (figure
4-47). This action is used to control the outboard engines for pitch,
yaw, and roll. Gimbal commands are accomplished by diverting
the flow of high pressure hydraulic fluids against either side of
the actuator pisto.n. The system consists of a servovalve, a filter,
a preflltration valve, a manually operated servovalve bypass valve,
an actuator position scale, a midstroke locking device, and valves
for fluid sampling and system bleeding. Two servoactuator as-semblies are mounted in a perpendicular plane that intersect at
the longitudinal axis of each engine, and are mounted between
stage superstructure and engine. The engines are moved, propor-
tionally to the magnitude of the electrical input signals, as theactuators extend or retract, independently or simultaneously. The
maximum gimbal angle is -4-8 deg; yet past operations have notused over one-third of this capability. The stroke of the actuator
is 9.560±.06 in., but is limited to 7.640±.06 in. To accomplish
1 deg of engine movement the piston moves 0.478 in. The operating
pressure on these equal-area pistons is 3200 psig. The servovalve
also completes the cycle by directing the return of low pressurefluid to the reservoir. The feed-back potentiometer sends a signal
to the flight control computer corresponding to the piston move-
ment, which is used to reduce the signal strength as desired actuator
CYLINDER
CYLINDER
BLI
HYDRAULIC
FILTER
!EDBACK
POTENTIOMETER
tECK
VENT
Figure 4-47
4-41
Section IV S-IB Stage
MONTORINGDEVICES
Ai
1C_
PRESSURE
INCREASE
Dt
CONNECTOR
©
FITTING
SUB-ASSY -_
CASE AND
MEMBRANE
ASSY ._ CONNECTOR-FLAN GE
ASSY
S
FLOW
PRESSURE TRANSDUCER
MOVABLE
CONTACT
ARM
ELEMENT
THERMAL SWITCH
SHAFT m INSULATION CASE
s"///////2/2/2//Z/2/2/2//"_/)_///////////////////A
.__ " ..... _,_%_,.._%_.'_'_'_'_\\"1B WHITE
_ A RED
C BLACK
D SPARE
FLUID LEVEL POTENTIOMETER C.H 14248-I
4-42 Figure 4-48
position is reached. The operations of the servoactuators are moni-
tored at the engine deflection panel and the analog recorders inthe LCC.
MONITORING DEVICES.
Various switches and indicators (figure 4-48) monitor the hydraulic
system operation within the accumulator-reservoir and manifold
assembly. High pressure relief valves protect the high pressure side
of the system by allowing high pressure fluid to vent into the low
pressure side. A low pressure relief valve protects the system frombeing overfilled. The pressure transducer, mounted on the manifold,
monitors high pressure in the hydraulic system. The pressure sensed
is converted to an electrical signal for the DDAS system and
checkout. The DDAS system provides a readout of this function
to a strip chart and a meter on the hydraulic panel in the LCC.
The capability range of pressure for the transducer is zero to 4000
psig. A thermal switch monitors the hydraulic fluid temperature
in the low pressure side of the system. When the fluid reaches
normal operating temperature 200±10 ° F the bimetallic sensing
element snap-actuates the contact arm and opens the circuit to
an indicating lamp in the LCC. When the temperature decreasesto 155+10 ° F the circuit will close and the indicator lamp will
light. The reservoir is equipped with a fluid level potentiometer.The movement of the reservoir piston is detected to give an equiva-
lent output voltage. The potentiometer is a single element,2000±100. ohm wire wound, linear translation-type unit and is
internally mounted in the base of the accumulator-reservoir. The
varying output voltage is processed into the telemetering system
to provide a direct readout and recording in the LCC.
HYDRAULIC HOSES AND TUBING.
The main pump and actuator flexible hoses have the same general
configuration but differ in detail construction. The outside coveris of stainless steel wire braid and an inner liner based on a teflon
compound. The low pressure hose designed operating pressure is
50 psig. The low pressure hose was tested hydrostatically at 3000
psig and performed without leaks for 5 min. The high pressurehose has an additional carbon steel reinforcement braid between
the inner and outer layers. The design operating pressure is 3200
psig. The high pressure hose was hydrostatically tested for 5 minwithout a leak at 6000 psig. The operating temperature range for
both high and low pressure hose is -65 ° F to 275 ° F.
The tube assemblies are made from seamless corrosion-resistant
steel, 304 or 304L controlled by MSFC-SPEC-131. The tubing is
flared in accordance with MC-146 and uses one sleeve and one
nut on either end of the assembly. The nuts and sleeves are of
corrosion-resistant steel. The nut used on high pressure applications
are in compliance with MF-818 for precision flared fitting ends
with interlocks.
ELECTRICAL.
Two independent bus networks (+ 1DI 1 and + ID21) distribute
primary power as shown in figure 4-49. Source power to these
buses originates from ground-base electrical support equipment
during prelaunch operations and transfers to stage batteries 50 sec
before liftoff. A separate battery supplies each bus network. Gener-
ally, one battery (DI0) powers operational systems that draw high
transient current and generate bus voltage transients. The other
battery (D20) powers the measuring system and supplies parallel
power for critical functions such as engine cutoff. This dual isolationfeature ensures more than adequate electrical power. Primary power
varies in the 27- to 30-Vdc range. Line loss is less than 2 Vdc
Section IV S-IB Stage
from stage bus to stage load. A total of 13 cables interconnect
the electrical system above and below the propellant tanks. Four
cables are in the cableway on fuel tank FI; nine cables are in
the cableway on fuel tank F2. Structure of the center lox tank
provides a unipotential ground path (IDCOM). All electrical equip-
ment incorporates design provisions in accordance with MIL-E-6051
(Electromagnetic Compatibility Requirements) to prevent electro-
magnetic interaction between electrical systems. Stage electrical
equipment operates independently of power sources from other
stages. However, some buses extend into the stage to form a feed-
back loop to indicate certain events. For example, the S-IVB stage
powers a bus (+ 4DI 1) that indicates physical separation of stages
and consequently initiates J-2 engine start logic; the S-IU stage
powers buses (+6D91, +6D92, and +6D93) that loop the H-1
engine thrust OK pressure switches to the emergency detection
system (EDS). There are other extra-stage buses, but none have
drain on the S-IB stage electrical system since the requesting stage
supplies them power.
BATTERIES.
Two batteries supply primary power to separate bus networks in
the stage electrical system; both batteries are identical and discus-sion is limited to one unit. The battery is a manually activated
storage unit containing 21 cells composed of silver oxide (pos) andzinc (neg) plates. The cell wiring arrangement permits selection
of either 18, 19, 20, or 21 cells so that battery voltage is between28 and 29.6 Vdc when measured on launch day under load condi-
tions. The battery is packaged dry (without electrolyte) to extend
shelf life; it is activated no longer than 168 hr before launch. Theactivation time was extended from 120 to 168 hr after test data
confirmed the extension had no significant effect on battery perfor-
mance. The activated battery is installed approximately 57 hr before
lifloff. Physical and electrical characteristics of the battery are
shown in figure 4-50.
Battery Construction.
Construction features protect the battery from damage and provide
for internal pressure control as illustrated in figure 4-51. The battery
box is composed of magnesium alloy for lightness and strength.It is coated for thermal emissivity effect and environmental protec-
tion. The cells are spaced with Neoprene shim-stock and pottedto form a unitized cell. block. A cover, molded of scotchcast and
filler composition, protects the top of the cell block. The box cover
adds support to the battery box unit and permits access to thecell block for inspection and cell activation. An o-ring gasket seals
the battery box and cover. The seal aids primarily in maintaining
the internal pressure as regulated by the pressure relief valve; it
also prevents entrance of moisture and other foreign matter.
Blind Plug Selection.
The blind plug assembly (figure 4-51) is a jumper connector that
completes the battery circuit when installed. The plug selects either18, 19, 20, or 21 ceils so that the battery output voltage (under
load) is 28 to 29.6 Vdc. Plug selection is based on actual load
test of the stage electrical system on launch day. A blind plugmade for the testload range is selected as shown in figure 4-52.
Initially, the blind plug is selected on the basis of the overall test
(OAT) bus current measurements before battery installation. After
installing the batteries (approximately 57 hr before liftoff), the stage
is powered up and the bus current is rechecked. If the load currentis the same as measured in OAT, the installed blind plug is used;
if not, the proper blind plug is selected.
Activities.
The battery is manually activated with potassium hydroxide (KOH)
4-43
Section IV S-IB Stage
ORKNO. I28VOC)ASUREMENTVOC)"--PRIMARYPOWERDISTRIBUTION
4-44
DISTR
+1D32
+ID34
+1D81
+1D82
Figure 4-4g
I •i •I •I •I •I •I •
IIImlmn IIm nm Dmuum mu j n RII_
n nim mmm mmmlunnmmimnn mnmlnmnNm nmnmimnm amHolll
+1DI0 DI0 BATTERY TERMINAL PWR +1D89 5 VDC MASTER MEAS
+IDll NETWORK NO. 1 PRIMARY PWR SUPPLY
÷1D12 NETWORK NO. 1 PWR TRANSFER J"_ENERGIZED BY ESE FOR PWR
BUS TRANSFER TO LV BATTERIES
+1D16 INHIBITS MULTIPLE ENGINE 50 SEC BEFORE LIFTOFF
CUTOFF_ENERGiZED BY MULTIPLE+1D17 PWRS TM LINK F1 (AND LIQLI D
LEVEL RACKS ON SA-206 ONLY) ENGINE CUTOFF ENABLE CMD
+ID20 D20 BATTERY TERMINAL PWR (SEE SEQUENCING)
+1D21 NETWORK NO. 2 PRIMARY PWR I_'_.ENERGIZED BY TOPS ENGINE
+1D22 NETWORK NO. 2 PWR _CUTOFF CMD (SEE SEQUENCING)TRANSFER BUS
+1D24 PWRS TM LINK PI AND MEAS _ENERGIZED BY ESE CMDRACKS (AND LIQUID LEVEL RACKS
ON SA-207, 208, AND 209) 5JSJSJSJSJSJS_ENERGIZEDAT PWRTRANSFER
SPECIAL NETWORK BUS (SW TO ARM +1D34
SEL CMDS)_ENERGIZED BY ESE AND
ENGINE TOPS C/O BUS LOCKED-UP BY STAGE BUS
5 VDC MEAS SUPPLY ÷1D22 AT PWR TRANSFER
5 VDC MEAS SUPPLY CH-14207.1
BATTERY CHARACTERISTICSDIMENSIONS
LENGTH: 18.0 INCHES
WIDTH: 7.5 INCHES
HEIGHT: 7.0 INCHES
WEIGHT
DEACTIVATED: 52 POUNDS
ACTIVATED: 58 POUNDS
ELECTRICAL
VOLTAGE (UNLOADED): 38.3 TO 39.3 VDC (2t CELLS)
VOLTAGE (LOADED)t 28TO 29.6VDCPOWER CAPACITY: 2000 AMP-MINUTES AT 100 AMPS
ELECTROLYTE
SOLUTION: POTASSIUM HYDROXIDE (KOH)
SPECIFIC GRAVITY: 1.40 + 0.042 AT 77_
VOLUME: 1890 ± 21 CM 3 P'ER BATTERY
DEGASSING PERIOD: 2 TO4 HOURS
PRESSURIZATION
BATTERY BOX MEDIA: DRY NITROGEN (GN2)
BATTERY BOX VENT: OPENS AT 26 PSID MAX;
CLOSES AT 10 PSID MIN
CELL VENT: OPENS AT 2 TO 6 PSID
TEMPERATURE
STORAGE: 68 TO 80°F
OPERATING: 68 TO 104°F
CH.14254-2
Figure 4-50
electrolytic solution having a specific gravity of 1.40-4-3 percent at
77 ° F. Each cell (figure 4-53) is activated with 90-----1 cc of solution
premeasured in individual containers. The battery chosen for flightuse cannot be older than 36 mo, nor can it be activated longer
than 168 hr before launch. There are two methods for adding
electrolyte to the battery, drip activation and vacuum activation.
Drip activation requires a rack that holds 21 individual activatorcases, each containing 90 cm _ of electrolyte. A drain needle, located
centrally within each case, facilitates insertion into the respectivecell filler holes. The rack is positioned to allow gravity feed into
the individual cells. Vacuum activation also requires a rack with
individual activator cases, but includes a vacuum pump as part
of the assembly. Outlet filler tips, one for each cell, extend fromthe lower side of the activator case. Filler adapter screws thread
into each cell filler port and connect flexible tubing to transfer
to electrolyte. The vacuum system consists of a pump, gage, regula-tor valve, and relief valve to evacuate and fill the battery cells.After activation, each cell vent valve (part of filler cap assembly)
is installed. The activated battery is subjected to a I0- to 20-A
load for 15 to 30 sec to verify proper activation and performance;
this test occurs 4 hr after installing the cell vent valves. The installed
vent valves allow internal cell pressure to vent at 2 to 6 psid but
prevent leakage of electrolyte regardless of battery position. Theactivated battery is stored at 68 to 80 ° F until ready for installation.
Battery installation and the last case isolation test occur no soonerthan 12 hr after installing the vent valves. Should the activated
storage time exceed 168 hr, the battery must be replaced.
Internal Shunt.
A shunt is the means for measuring battery current during flight.
Figure 4-52 shows the shunt connected in series with the negative
lead of the battery output; monitoring connections are available
to the telemetry system. The shunt design allows for an output
voltage of 100-+3 mVdc between pins A and B 02) when the battery
is delivering 75-A current at 77 ° F. The voltage drop across the
shunt is proportional to battery load current.
Thermistor.
A thermistor in each battery (figure 4-52) provides battery temper-ature measurements. A calibration curve for this measurement is
Section IV S-IB Stage
provided with each battery with a _ 1% accuracy. Battery temper-atures are continuously monitored after installation in the stage.
An increase in battery internal temperature indicates that a cell
failure is about to occur.
Predicted Performance.
Data from previous flights indicate that the stage power system
is both adequate and reliable. For example, battery current
averaged only 20A during the SA-205 flight. The D10 battery used
7.7% of the rated capacity while the D20 battery used 7.3 percent.
Voltage output averaged 28 Vdc during this flight. A summation
of the estimated voltage and current requirements for The SA-206
flight is profiled in figures 4-54 through 4-57, respectively. Analysisof the electrical load requirements for the flight shows that even
a faulty current up to 60 A from either or both batteries would
not hamper critical systems operation. Complete current loss from
one battery would neither prematurely cut off the engines nor
impair normal engine operation.
MEASURING SUPPLIES.
Measurements require voltage in the 0 to 5 Vdc range; three
measuring supplies provide regulated voltage to instrumentation
Notes
POWER CABLES ARECELLS ARE SPACED _ CONNECTED, CHECKED,WITH SHIM
STOCK AND POTTED TO AND POTTED INTO CABLE
FORM CELL BLOCK. WELL TO UNITIZECELL BLOCK.
[_ COVER IS CASTDIRECTLY ONTO CELL
' "__a-----.....,,__ BLOCK AFTER INSTALLING
- _ PLUGS TO FORM CELL_ ACCESS HOLES WHEN
___._ MATERIAL CURES.
_' __ _J "_---BATTERY BOX
• .;o,.. ___ (O-R,NO);if. . .._i .° .P.. o.:.o: 0 ° ..'..
• - ...: ..._ :...-:_. :.
, __ _- CELLBLOCKCOVER '_]
___/_ ,,___ I (SCOTCHCAST/F,LLER)
CELt• BLOCK
__/ (21CELLS)_>
_ CABLE WELLg>
,_.D ELL_L-_TOR(_)
%----PWR. OUTPUT (JI)
GH 14253-1
Figure 4-51
4-45
SectionIVS-IB Stage
BATTERY WIRING DIAGRAM CELLCONSTRUCTIONDIO AND D20 BATTERY _ --
__5(_)6_+7_1011 12 13 14 15 16 17 _1+D_[I', 1
I ° I + I- I +1 - I + i- I + I
SHUNT
J2 J1
TM _
OUTPUT OUTPUT
Nolo
J3 CONNECTOR MATES A BLIND PLUG
THAT SHORTS TERMINALS AS SHOWN TOGIVE 28 TO 29.5 VOC UNDER LOAD AT JI.
SEE TEXT (BLIND PLUG SELECTION).
_ IILINB rI.Ug _----_
BUS NO. PINS
CURRENT CELLS SHORTED
< 3A 18 FD
3-55A 19 ED
SS- 100A 20 BD/CD
• 100A 21 AD
CH.14249.1
Figure 4-52
tbr this purpose. Known as the D81, D82, and D89 measuring
supplies, the units each supply independent networks. Unit D89
powers instrumentation located above the firewall; the other two
units (D81 and D82) provide power to instrumentation located
below the firewall. Each unit receives 28-Vdc input from an as-
sociated measuring distributor. Solid-state circuitry converts this
input to 5 Vdc_0.25 percent. This precise voltage returns to each
measuring distributor for subsequent distribution to instrumenta-tion.
DISTRIBUTION.
Seven distributors route electrical power (see figure 4-49); one
power distributor, one main distributor, two propulsion distributors,
and three measuring distributors. Cables transmit primary power
from the batteries to the power distributor, establishing the two
primary network buses. From this origin, secondary buses are
established in the other distributors located in strategic areas (see
Equipment Location) above and below the propellant tanks. Sub-
systems below the tanks interconnect through two propulsion dis-
tributors and the two remaining measuring distributors. Cabling
reaches the distributors through two protective cableways; one onfuel tank FI, the other on fuel tank F2.
Power Distributor.
The power distributor routes battery power to auxiliary distributors
4-46
CELL FILLER
RETAINER
CELL
OXIDE
PLATE
PLATE
Neto
EACH CELL IS ACTIVATED WITH
90cc POTASSIUM HYDROXIDE
(KOH) ELECTROLYTE.
OH 14250-I
Figure 4-53
that act as junction boxes and switching points. Battery terminal
buses +ID10 and +lD20 (also ESE buses +AIDI and +A2DI)
are powered in this distributor. At 50 sec before liftoff, a command
for power transfer from ground sources to stage batteries closes
heavy relay contactors and establishes the two primary buses
+ IDI 1 and + ID21. Two pyrotechnic switch assemblies attached
to the distributor ensure this transfer shortly after liftoff (see Power
Transfer).
Main Distributor.
The main distributor supplies power mostly on command from
the switch selector (see Sequencing) and contains relay circuitry
that switches + 1DI 1 and + ID21 power to operate or enable flight
control components. It provides interface control for normal 4-by-4
engine cutoff (see H-I Engine Cutoff) and EDS engines shutdown
(see Section VI). Distributor circuitry inhibits multiple engine cutoff
should one engine shut down between 3.0 and 10 sec of flight.
Circuitry also prevents premature shutdown of one OB engine from
affecting lox depletion cutoff. By enabling the propellant level
sensors, it permits the separation EBW units to charge when onesensor actuates. The distributor triggers the EBW units on stage
separation command. Since the main distributor performs manycritical functions, its reliability has been closely checked. The most
critical components in the unit are the electromechanical relays.Tests to determine contact chatter (open and close periods lasting
100/_sec) and the effect of intermittent output show that the distrib-
utor is both reliable and qualified to perform flight sequence andEDS functions.
Propulsion Distributor.
The two propulsion distributors contain circuitry, to detonate the
Conax valve squibs and subsequently cut off the engines upon
command (see H-I Engine Cutoff Commands). The distributors
also operate the engine prevalve control valves. One distributor(9AI) controls propulsion functions on the OB engines, the other
distributor (9A2) controls the IB engines. Prior to SA-202, only
one propulsion distributor was used. Addition of the second unit
Section IV S-IB Stage
31.0
0> 30.0
Z
_ 29.0
28.0
27.0 ..,_,_ _ | I I t-253O 0 10 20
' _C_OFF__ --\ $I
I NBOARD _-- CUTOFFENGINES
OUTBOARD
ENGINES
'Jl,'_ Jo!1,_o ' ' ' '_¢I. ' '180 200135 140 145 1503O
FLIGHT TIME (SEC)
CH-20061
Figure 4-54
32.0 "_"-'_ _-"r-"_-$ _, i I I I _ I i It'$ I ] I I (_1 i i
31.0
g
o>._ 3o.oz
_ 29,0
N
28.0
_.o .--......_...J..----.4_ I I I-25_ ;-_# 0 10 20
CUTOFF,,IINBOARDENGINES _ CUTOFF
OUTBOARD
ENGINES
• = I I I| I I I I _11 1_ 23O3 _ ; 1_ 1_ 130 135 1_ 145 1_
FLIGh3" TiME (SEC)
CH.2006
Figure 4-55
4-47
SectionIV S-IB Stage
--ased reliability r ,d provides circuitry for a third thrust OK
pressure switch on each engine. Each distributor now has 2-of-3
voting logic for thrust decay cutoff on each engine.
Qualification tests on the distributor showed some relay chatterin excess of 100 /_sec duration, but the maximum duration did
not impair circuit operation. Also, some wire strands broke duringvibration but none of the effected circuits failed; flight distributors
now have improved wiring arrangement.
Measuring Distributors.
Three measuring distributors route 28 Vdc to power signal condi-
tioning equipment and 5 Vdc to energize measuring devices; the
5-Vdc source is also a reference and calibration signal for measure-
ments and telemetry (see Instrumentation). Each distributor powers
a 5-Vdc power supply (see Measuring Supplies) that returns the
converted and regulated 5-Vdc source. The distributors supply three
independent networks. One distributor (12A26), located above the
propellant tanks, services the D89 network for measurementsthroughout the entire stage. The other two distributors (9A3 and
9A4), located below the tanks, service the D81 and D82 networks
for engine measurements. With this arrangement, loss of one ne-
twork would not impair critical measurements. Vibration tests havequalified all three distributors for flight use.
SEQUENCING.
A stored program in the launch vehicle digital computer controls
inflight electrical sequencing; a switch selector is the stage com-
munications link. The switch selector decodes incoming computer
signals and activates the proper circuits in the main distributor.Figure 4-58 illustrates the results of the switch selector commands.
These commands are sequenced in four time bases (T°). A specific
milestone event establishes each time base. If any one time base
is not established, subsequent time bases cannot be started andthe vehicle cannot continue the mission. Likewise, programmed
safeguards prevent premature initiation of a time base. See Section
VI (Navigation, Guidance, and Control) for a closer examination
of program events sequenced from the LVDC and a definitionof the time bases.
Switch Selector.
Each stage has a switch selector for communicating with the guid-
ance and control system in the S-IU. Functionally, the switch
selector receives coded commands from the LVDC (via the LVDA),
decodes them, and activates the selected circuits. Use of a coded
message decreases the number of interface lines and increases
programming flexibility. The switch selector is divided into two
functional sections to maintain power isolation between stages: the
input section (relay circuits) receives power from the S-IU stage;
and, the output section (decoding circuitry and drivers) receives
power from the selected stage. The input and output sections couple
through a diode matrix. This matrix decodes the 8-bit command
input code and activates a PNP output driver, thus producing the
switch selector command to stage system. The switch selector exe-
cutes commands given by the 8-bit code or by its complement.
The switch selector operates on positive logic, that is, +28 Vdc
for a binary 1 and 0 Vdc for a binary 0.
A flight sequence cohmand from the LVDA loads the switch selector
register with a coded word command; the word format is shown
in figure 4-59. Register bits 1 through 8 represent the functional
sequence command. Bits 9 through 13 key the selected stage. Bit
14 resets all the relays in the switch selector when the LVDC receives
faulty verification information. Bit 15 activates the addressed switch
selector to read the command and produce the proper output.
The LVDC loads the switch selector register in two passes; bits
1 through 13 load during the first pass and, depending on the
ESTIMATEDDIO BATTERYLOAD PROFILE
30 --
20.
1°- 5
CUTOFF
INBOARD 38.0
ENGINES---_
18.3
LIFTOFF -_ \
15, " " " " "
I _ POWER TRANSFER
42.4
CUTOFF
OUTBOARDENGINES
SEPARATION
28.427.8
1 I I I i; I i i :| I I I I i|11 I I0 10 20 30 110 120 130 135 140 145 150 180 200
TIME AFTER LIFTOFF (SEC)
CH-2006:
4-48
Figure 4-56
feedback code, either bit 14 or 15 loads during the second pass.
After setting the 8-bit command, the received code complementis sent back to the LVDC through eight parallel lines. This feedback
(verification) is returned to the digital input multiplexer of the
LVDC and is subsequently compared with the original code in
the LVDC. If the feedback agrees with the original code, a read
command is given. If the feedback does not agree, a reset command
is given (forced reset), and the LVDC reissues the 8-bit command
in complement form.
POWER TRANSFER.
Primary power transfers from ML electrical supplies to internal
stage batteries at 50 sec before liftoff. The command issues fromthe terminal countdown sequencer to the stage power distributor
(12A25) at this time; however, relay logic in an ML programdistributor restrains the command until confirmation signals in-
dicate that the S-IB, S-IVB, and S-IU stages are ready for transfer.
The command for transfer closes relay contacts connected directly
between the stage batteries and network buses in the power distrib-
utor as shown in figure 4-60. Parallel transfer circuitry is providedwhen liftoff movement detonates two squib-actuated switches that
complete a backup circuit. Gas pressure generated by the exploding
bridgewire (see Ordnance) actuates ganged contacts in the switches
that parallel contacts on the transfer relays (KI and K2). Theswitches also maintain power on these relay coils to nullify contact
chatter. When the switch assemblies are installed (1 day, 13 hr
before launch) on the power distributor, a series circuit completes
through an electrical loop in each switch and enables a safety switch
installed signal to a display in the LCC (S-IB networks panel).
MEASUREMENTS.
Measuring instrumentation monitors switch selector operation,measurement bus voltage, primary network voltage, and battery
Section IV S-IB Stage
current. These measurements are listed in figure 4-61, which shows
the display ground station for each measurement.
HEATERS.
Cryogenic characteristics of the propellants (see Propulsion) requirethat certain components be heated to maintain operable temper-
atures. These components, which either immerse in the propellants
or operate in the transfer system, are as follows:
a. Turbo Pump Bearing No. 1
b. Fuel Additive Blender Unit Gearcase
c. Main Lox Valves
d. Lox Vent Valves
e. Lox Overfill Sensor
f. Lox Level Sensors
g. Fuel Level Sensors
Each component has an integral resistance-type heater that is
thermostatically controlled. The heaters receive power through a
115-V, 60-Hz, 3-phase, 4-wire system from ESE. Temperatures of
the components are sensed by onboard thermistors. The resultant
temperature indications are made available to the ground computerfor evaluation. The heaters are turned on prior to loading cryogenic
propellants approximately 8 hr before liftoff and are turned offat the ALL ENGINES RUNNING indication.
EQUIPMENT LOCATION.
Electrical power assemblies and instrumentation equipment arelocated in the upper and lower skirts of fuel tanks FI, F2, F3,
and F4 as shown in figure 4-62. The upper portions of tanks FI
and F2 have extended skirts that house most of the equipment.
40 --
30
20
10
LIFTOFF --_ 22.4
20.4 _|
_- POWER
TRANSFER
I I I_'_ ( o 1o 20
CUTOFF I NBOARD
ENGINES_ 38 2 38.4
i I rcUToFE°UT_A"°
22.6 24.2 J _ 24.1 ts
/ I I, _ I I _l' I I i I '_I 18o 20oi 110 120 130 135 140 145 150
TIME AFTER DFTOFF (SEC)
CH-2_&
Figure 4-57
4-49
SectionIV S-IB Stage
ELECTRICAL SE¢ UENCING
,1F
Hales
THIS CIRCUITRY IS SIMPLIFIED TO SHOW ONLY THE
RELAYS AND DIODES NECESSARY TO EXPLAIN THE
RESULTS OF THE SWITCH SELECTOR COMMANDS.
ENGINE 1 CUTOFF RELAY CIRCUIT IS TYPICAL OF ALLOUTBOARD ENGINES. TOPS RELAY VOTING CIRCUIT IS
TYPICAL OF ALL 8 ENGINES. ENGINE 5 CUTOFF RELAY
CI RCUIT IS TYPICAL OF ALL INBOARD ENGI NES.
RELAY K66 CONTACTS ARE IN OUTBOARD ENGI NE
CI RCUITS ONLY.
[_ ONLY ONE OF THE TWO EBW FIRING UNITS
IS SHOWN FOR CLARITY. BUS +1D21 POWERS
THE NO. 2 UNIT.
[_ ACTUATION OF THE PROPELLANT LEVELSENSORS STARTS TIME BASE NO. 2. THERE
IS NO SWITCH SELECTOR COMMAND.
_ NORMALLY O.B. ENGINES CUTOFF STARTSBY SIGNAL OF PROPELLANT DEPLETION.
HOWEVER, TIME BASE NO. 3 STARTS WITH A
COMMAND FOR CUTOFF WHICH IS ALSO A
SAFEGUARD AGAINST STAGE SEPARATION
WITH THRUST PRESENT.
[_ BUS +1D21 PROVIDESREDUNDANT CHARGI NG
CMD THRU PEDUNDANT
FUEL & LOX LEVEL SENSORS
(9A69 & 9A70 NOT SHOWN)
FOR SEPARATION & RETRO-
ROCKET EBW FIRING UNIT
SYSTEMS NO. 2.
_IISEQUENCE "'_ RECEIPT,"
immmmumm_,
SWITCH
SELECTOR
CHAN. NO.
100
16
15
2
39
2
39
I04
98
97
79
18
20
TIME
BASE
(SEC)
T 1 + 3.0
T 1 + 10.0
T 1 + 10.1
T I _ 20.0
T 1 + 25.0
T 1 + 119.8
T I + 124.8
T 1 + 129.5
T2 + 00.0
i T2 + 3.0
IT 2 +4.5
T2 + 5.0
T3 + 00.0
T 3 +0.1
T 3 + 1.3
4-50
S-[B/S-_VBSEPERATION CMD.
+IDll ÷lOll +IDll
--fK15 -T CHARGE
TRIGGER+IDCOM
SEPARATION& RETRO ROCKET
EBWFIRI NGUNIT SYSTEMS
NO. 1; 11A13, 11A26, 11A28,
11A30, & 11A44
1DCOM
PROP. LEVEL8 SENSORS ENABLE
÷IDll +ID21
_ K47 _K22 TK22
+IDII
FUEL LOX
LEVEL LEVEL
SENSOR SENSOR
9A63 9A68
MAIN
DISTR
12A1
PROPULSION
K75 K_] DISTR. 9AI
,c:_b k+1012
+IDCOM +lD12
Figure 4-58 (Sheet 1 of 2)
+1D11
+IDCOM
K43 •
MAiN
DISTR.IDCOM 12AI
TM CALIB. ASSY.
POWER SUPPLY
13A25
PLUG
K2 TYPE
J-BOX
13A26
TMCALIB. ASS'Y
13A495
CH .20065
Section IV S-IB Stage
MAIN DISTR. 12A1
+ID32 +ID21 +ID11
+IDI2
I_C K - - IDCOM
I
1DCOM
O O O Kll ENGINE 1
; _ ,_ CUTOFF RELAY(TYP. OF ENG.
"_" _ _ 2, 3, & 4)(6 (6 0Z Z Z
2 0 0
PROPULSION SYS. DISTR.
9AI
o
1DCOM
K31_
K70 &K70 j
1DCOM K41_
+IDI2
+1Dll
+1Dll |
1DCC_V_ +ID21
i T T]I+1Dti
-K4 Y K22 1DCOM"A--LIFfOFF T SEE PROP
i LEVEL SENSORS
l l ENABLE
K15 ENGINE 5 0CUTOFF RELAY
(TYP. OF ENG. "_
6, 7, & 8) 0
I!>PROPULSION SYS. DISTR.
9A 2
IK17- =
C-H 20066
O O O
0 (6 0Z z Z
+ ID16
: _-Klli
s_K41
K31
TOP 1
2 OF 3:OTING _K51
1 [ IDCOM
K1 +T PWR ID34
+1Oll _l[ XFR _KI'"
Y PWR i!2; 0+1Dll K18_XF R &K
',10 10
+ IDll + ID21 IDCOM
+1Dll
Figure 4-58 (Sheet 2 of 2). 4-51
Section IV S-IB Stage
A bulkhead seals this area above each tank and forms an air-tight
compartment that is conditioned (see Environment Conditioning)
for increased reliability of electrical components. Compartment no.
1 (FI) contains mostly electrical power equipment while compart-
ment no. 2 (F2) contains mostly telemetry and tracking equipment.
The lower portions of all tanks are open and house propulsion
and measuring equipment. Electrical cabling reaches the upper and
lower tank areas through two cableways, one on tank F 1 and theother on tank F2.
S-IB POWER TRANSFER
INSTRUMENTATION.
The S-IB stage instrumentation system meet the following objec-tives:
a. Provide in flight data on critical subsystem performance
b. Provide prelaunch ground monitoring of stage subsystem opera-tion
Measuring and telemetry subsystems comprise the S-IB stage in-
strumentation. The major instrumentation equipment is located onthe interior of the forward and aft skirts of fuel tank F 1, the forward
and aft skirts of fuel tank F2 and the aft skirt of fuel tanks F3
and F4 (figure 4-62).
MEASURING SYSTEM.
The measuring system (figure 4-63) consists of transducers, signal
conditioners located in the measuring racks, and measuring distrib-
utors. A total of 283 measurements are made on the S-IB stage
of vehicle AS-206; 266 of these are telemetered to ground stations
during flight. Prior to launch, 52 measurements not used in flight
are transmitted to the block house by telemetry or hardwire connec-tions.
Measurement Numbers.
A coding system is employed to provide identification for each
measurement. The coding system used for measurement numbers
describes the particular type of measurement and the area location
on the S-IB stage. A typical measurement number is explainedin figure 4-64.
Signal Conditioning.
Approximately 86 of the source signals in the measuring program
are unsuitable for use by the telemetry system; therefore, signalconditioning is required to modify the measurements before deliv-
ery to the vehicle measuring distributors. Replaceable plug-in
signal conditioning modules are installed in the measuring racks.
,WITCH SELECTOR
REGISTER WORD FORMAT
LVDA LOADS
REGISTER WITH
BINARY WORD _ MODE
+ 28 Vdc IS BINARY I _¢ STAGE SELECT r--T--
0V_ISBINARY0 _ I I
o'11='
0 I 0 0 0 I 0 0 1 0 0 0 0 0 1
BINARY CODE FOR
S-IB INBOARD
ENG CUTOFF
Figure 4-59
C.H 14433
4-52
MI
J+lOll0 PROGRAM DISTR202- 21AI
_,e READY IIs-, BREADYITS-IBREADYI
2 'IDIR VEHICL[
PWR
EXTERNAL
I
12A39
+1010
POWER DISTR
L-- 12A25
+1011
j SWITCHASSEMBLY J
I '2A25II I
L._ __ __l
BUSES TO POWER BUS +IDI2 & +1D22
THIS BUS POWERS TRANSFER
RELAYS (KI) IN OTHER ONBOARDEQUIPMENT.
J MAI N DISTR J12A1
I II +lb. I
SWITCHASSEMBLY'S
.-K14 ARM j
"K4 LIFTOFF j
II
t
F INSTRUMENT
COMPARTMENT
NO. 2 (FUEL TANK F-2)
PANEL
I ASSEMB -- 227_222-_-2E-_-_-_-_-_-27-2-__
{12A25)_
SWITCH _
(12A32)
Figure 4-60
CH-14512.2
MEASUREMENT
NO. J TITLE
VK1-12 ] SW SEL COUNT IND
VKI15-12 I SWSELREGTEST
VM1-9 I MEAS VOLTAGE +1D81
VM2-9 [ MEAS VOLTAGE +1D82
VM9-12 J MEAS VOLTAGE +1D89
M16-12 J D21 BUS VOL:[AGE
M17-12 I
XM18-12 I
XM19-12 q
i
D 11 BUS VOLTAGE
D10 BAT CURRENT
D20 BAT CURRENT
ESE
X
X
X
X
X
Figure 4-61
DISPLAy
AUX MCC
X
X
i xI
i X X I
X X
iC-H 14432
Six measuring racks are located in the S-IB stage. One rack islocated in the aft skirt of each of the four fuel tanks and two
racks are located in instrument compartment no. 1. The conditioned
outputs of the measuring racks are sent to measuring distributors.
The S-IB stage contains three measuring distributors. Measuringdistributor 9A4, located in the aft skirt of the F2 tank, controls
direct measurements and routes the outputs of measuring racks9A520 and 9A526, located in the aft skirt of tanks F2 and F3,
to the proper telemetering system. Measuring distributor 9A3,located in the aft skirt of the Fl tank, controls direct measurements
and routes the outputs of measuring racks 9A516 and 9A530,
located in the aft skirt of tanks Fl and F4, to the proper telemetering
system. Measuring distributor 12A26 located in instrument com-
partment no. 2 routes information signals and the outputs of mea-
suring racks 12A439 and 12A440 to the proper telemetering system.
Checkout.
A remote automatic system permits a remote checkout of the flight
instrumentation system. Each of the signal conditioning modules
contains a printed circuit board that includes transducer simulationcircuits for calibration testing of the module. Two relays are incor-
porated into the module: one relay for the upper (HI) end of the
calibrated range of the measurement and one for the lower (LO)end. The run mode returns the measurement to its normal operating
position. A control panel in the block house enables personnelto select the desired measurement module in the vehicle and the
calibration mode (HI, LO, RUN). Any number of channels can
be selected and energized in any of the three modes (HI, LO, RUN)
individually, or in sequence. During simulated flight tests or launch
countdown, a previously programmed computer may automatically
operate the system.
A measuring rack selector (9A546) located in the aft skirt of tank
F2 is used to decode measuring rack selection commands fromthe ESE, select the rack or racks to be calibrated, and pass on
ESE calibration commands to the selected rack or racks. The mea-
suring rack selector is used only during prelaunch checkout. Figure4-65 lists the test specifications.
TELEMETRY SYSTEMS.
Measurements made on the S-IB stage are transmitted over two
telemetry systems, GP1 and GFI (figure 4-66). Components of
the telemetry system are located in instrument compartment no.
1 (see figure 4-62) except the remote digital submultiplexer 9A700which is located in the aft skirt of fuel tank F2. The two telemetry
systems transmit data through a common antenna system consisting
of multicoupler 13A494, power divider 13A470, and antennas
Section IV S-IB Stage
11A494 and 1 IA495. The antennas are installed on opposite sides
of the stage at positions II and IV. A telemetry system calibrator,
13A495, supplies calibration voltages to the telemetry systems for
prelaunch checkout and inflight calibration. The major componentsused in the telemetry systems have been qualified in previous flights.
The momentary data drop-outs that have occurred in the past have
been traced to antenna pattern null points resulting in insufficient
signal strength, or ground station problems.
TELEMETRY SYSTEM GP1.
Telemetry system GPI is a PCM/DDAS link that transmits real-time checkout data before launch, and measuring program infor-
mation during flight. The telemetry system consists of Model 270
multiplexers 13A440 and 13A484, remote digital submultiplexer9A700, Model 301 PCM/DDAS assembly 13A485, and RF assem-
bly 13A486. The pulse-amplitude modulated (PAM) data from the
Model 270 multiplexers are converted to digital words in the Model
301 assembly. The Model 301 assembly then encodes the digital
words with a digital word from the remote digital submultiplexerinto a data frame. The data frame is transferred to the RF assembly
where it is converted to an RF signal. The RF signal is connected
to the RF coupler for transmission to the antenna system. Thedata frame is also used to modulate a 600 kHz, voltage-controlled
oscillator (VCO) the output of which is connected to the LCC
through a coaxial cable for prelaunch checkout and monitoring.
REMOTE DIGITAL SUBMULTIPLEXER.
The RDSM provides additional data-handling capability to the
pulse code modulation (PCM) telemetry system. The RDSM can
accept a maximum of one hundred inputs, which are sampled
sequentially in groups of ten. The ten outputs of the RDSM,
containing the information provided by the 100 inputs, can there-fore furnish the PCM telemetry system ten digital words made
up of ten bits each. The assembly handles digital information only.
Its inputs and outputs are set voltage levels that represent ON-OFF
conditions or binary numbers. The sequential sampling action of
the RDSM is controlled by timing signals from the Model 301
PCM/DDAS assembly.
MODEL 270 MULTIPLEXER.
The Model 270 Multiplexer consists of 23 submultiplexers, each
of which gate ten data sources into a 30-channel main multiplexer.Four data channels are continuously supplied to the main mul-
tiplexer; the remaining three channels convey frame identification
and synchronization pulses. The data frame output of the main
multiplexer is supplied to a PAM amplifier through a gate. The
PAM amplifier amplitude modulates a train of pulses in accordance
with the sequential data input. The PAM pulse train is then supplied
to the Model 301 assembly. The output gate, after the main mul-
tiplexer, allows a calibration signal to be substituted for data whencalibration of the data channels is desired. An internal calibration
generator produces a sequence of five calibrated voltages upon
• receipt of a calibrate command. Voltage levels of 0, 25, 50, 75,
and 100 percent of maximum level are sustained for ten frames
commencing with the master frame following receipt of the com-mand.
MODEL 301 PCM/DDAS ASSEMBLY.
The PCM/DDAS scans the pulse amplitude modulation wavetrains
from theModel 270 multiplexers in a programmed sequence and
multiplexes these wavetrains into an analog-to-digital converter.
The PAM samples in the wavetrain are encoded into a 10-bit digital
word and stored in a 10-bit register. An output gate multiplexes
the data in the register with data in the register of the remote
digital submultiplexer. The serial data words are then supplied
4-53
Section IV S-IB Stage
EQUIPMENT LOCATION
NoteCONDITIONED INSTRUMENT
COMPARTMENT, TYPICAL OF FUEL TANKS F3 AND F4_ FUEL TANKS FI (IC NO. 1) ARE NOT SHOWN, MEASURING
_ A,oP2<,c,o. 2> RACK,AS26,S,NA. SK,RT!_-_!_!!::."."ii;i OF TANK F3, RACK 9A530 IS
::!iiii:_"?-_ ql_ IN AFT SKIRT OF TANK F4
;i;i:i:_:!:!_ TM MULTIPLEXER
i:i:i:!:i:!:i:_ RECEIVERS \ 13A450--_ "_ 13A494"1 t-"TM CALIBRATOR
!i:Z!i::iii::i_ 13A499 __ _ ____| / -- T3A495
_ _ _,_,_%o;_¢,_,K ,O%OUPLER_?A___\ \ / "I%--TM"WRD'V'DER"_' FI, F2 F3 ANDF4 1_, , ,_A_0_--.A\ \._----_,-_---_M------L_ I I_ _'°
TANK PWR DIVID R
D81 MFJ_SURING ' " _ ,_'_"I_IIPOWER SUPPLY/ _ LY,A_ __ER\ _ _/_/I;._2_
...,,,,--" II i I, _-.-_ _ _,Cs'y_°,AS
i %_ Ii _LUL.I.U__ ",_t_ I", _3A,_ ,'4
\..,/1_1 ,I_ / /}-,J_, "_i __-- -_/×/ _ , % _-
k., _%3G--_.._--X" ' ' _ COMEARTMENTNO.,", _ _--_--_ \ _ FWD FUEL TANK FI /.--\.-
_ _ _ _ _/ ._ _ _ MEASURI NG BATTERY ASSY
I __ __- _ DISTRIBUTOR PWR DISTRIBUTOR 12A33
Z.-M_SUR_N_ I -- ,A3 ACCEL. lzc2s_--X---_- _'-_
FUEL TANK F1 /" \ _ I \ _',
M_S.D_STR;EU'rOR\_ I----X_--_--__L\ \
L, .--'_i.ll Ill -_._, :: I',_T'_'_'_, S_U_PL%_NG':i'_'I_ I I '
'\,, _\_---___ .._'_" ' ..' LM_R,NORACI! L-_,TC.SE,EaOR,_,,,,
,A_,0 APTSK,RT COM,ARTMENTNO.,FUELTANK" PWO_UELTANK_,
4-54Figure 4-62
to the 600 kHz VCO and the RF assembly. The 600 kHz FM
carrier comprises the DDAS output which is connected, througha coaxial cable, to the DDAS receiving equipment in the LCC.
This signal is demodulated, demultiplexed, and decoded, to recreate
the original measurements. These measurements can be displayedon meters, or recorded, or used as command verifications in the
automatic checkout equipment.
RF ASSEMBLY--TELEMETRY LINK GP1.
The RF assembly consists of a filter, a frequency shift keyer, and
a power amplifier. The digital data pulses are first filtered to reducesideband components. The filtered pulses then modulate the fre-
quency shift keyer. The FM signal is then amplified by the power
Section IV S-IB Stage
amplifier and supplied to the RF coupler as a carrier of 256.2MHz at a nominal 15 W with a modulation deviation of +36
kHz. The data rate is 72 Kbps.
TELEMETRY SYSTEM GF1.Telemetry System GFI is a FM/FM link (frequency modula-
tion/frequency modulation) which transmits measuring program
information during flight. The telemetry system consists of Model
C1-18 assembly 13A450 and RF assembly 13A451. The GFI
telemetry system is provided to convert analog measurement signals
into proportional frequency-intelligent signals for subsequent mod-
ulation of an FM transmitter. All input signals must be precondi-
tioned to a 0-Vdc to 5-Vdc range.
MEASURING SYSTEM
TRANSDUCERS OR TYPE MODULE
SIGNAL SOURCES USED
DC AMPLIFIERTEMPERATURE, PRESSURE,
STRAIN, CURRENT
VIBRATION
ENGINE CUTOFF SIGNALS
LIQUID LEVEL SIGNALS
ACCELERATION
RPM
PULSE DETECTOR UNIT
DISCRETE LIQUID LEVEL RACK
(SEPARATE PACKAGE)
SERVO ACCELEROMETER UNIT
FREQUENCY DIVIDER
Note
SIGNAL CONDITIONING
MODULES ARE CONTAINED
IN SEVEN MEASURING RACKS.
THE RACKS, MODULES, AND
TRANSDUCERS SHOWN ARE
TYPICAL.
MEASURING RACK
SIGNAL CONDITIONING
MODULES
SIGNALS NOT REQUIRING CONDITIONING FEED DIRECTLY
Figure 4-63
CH-14252-2
4-55
Section IV S-IB Stage
TELEMETRY COMPONENTS.
Model C1-18 Assembly.
The Model CI-18 assembly consists of 17 subcarrier oscillators
operating at standard inter-range instrumentation group chan-
nels 2-18. Channel 1 is not used because the power supply of the
guidance system in the IU operates at the same frequency. Thesubcarrier oscillators (SCO) are the units within the telemeter that
convert the analog inputs into frequency-intelligent data. Each SCO
is a voltage controlled oscillator that is set at a precise frequency
range. The frequency output will shift up to the high limit as the
input signal increases to 5 Vdc and shift down to the low limit
as the input decreases to 0 Vdc. Thus, the output frequency of
each SCO is set to be linearly proportional to an input of 0 to
5 Vdc. A gate at the input to the subcarrier oscillators allows a
calibrating signal to be substituted for the data. Calibrating com-
mands and voltages from the telemetry system calibrator are sup-
plied through the program plug assembly to actuate the gate and
check the operation of the channel. The program plug assembly
is a wiring option that allows a channel to be preflight calibrated
only, inflight calibrated only, or not calibrated at all. The plug
is wired in accordance with the measuring program for the particu-
lar vehicle• The FM outputs of the SCO's are mixed in the mixer-
amplifier and the composite signal is supplied to the RF assembly.
RF Assembly--Telemetry Link GF1.
The RF assembly uses the input signal to frequency modulate a
MEASUREMENT NUMBER EFINITION
VJ ESE DISPLAYHARDWIRE
AUXILIARY DISPLAY
PREFIX
A ACCELERATION
B i ACOUSTICS
C TEMPERATURE
D PRESSURE
E VIBRATION
F FLOW
G POSITION
H GUIDANCE &
CONTROL
VX D 29 o 1
PARAMETER
J RF & TELEMETERING
K SIGNALS
L LIQUID LEVEL
M VOLTAGE, CURRENT &
FREQUENCY
N MISCELLANEOUS
,R ANGULAR VELOCITY
J_ RPMSTRAIN
C-H 142_
Figure 4-64
KSC _RELAUNCHTEST SPECIFICATIONS•
KSC PRELAUNCH TEST SPECIFICATIONS
MEASUREMENT
ALL EXCEPT THOSE LISTED BELOW
DI-1 THROUGH -8
DISCRETE LIQUID LEVEL
EVENTS
SPECIFICATION
OUTPUT SHOULD BE PREDICTED
VALUE +2% FULL SCALE.
OUTPUT SHOULD BE PREDICTED
VALUE _+1% FULL SCALE
SIGNAL LEVEL SHOULD BE PREDICTED
VALUE 4- 10% OF LEVEL SPECIFIED
CHECK FOR GO=NO-GO OPERATION
4-56
Figure 4-65
C-H 14259-1
VHF oscillator in the FM transmitter. The 1.5-W signal is raised
to 20 W nominal by a power amplifier• The high power signal •
is then filtered to suppress harmonics and supplied to the telemetry
system RF coupler.
RF Multlcoupler.
The multicoupler uses tuned cavities to selectively pass the signals
to be coupled. Each carrier input passes through a resonant cavity
that is tuned to 1/4 wavelength of the carrier center frequency.
The coupler provides 18-db isolation between adjacent frequencies•
The carriers are then capacitively coupled together and routed tothe power divider.
Power Divider.
The power divider equally divides the composite carrier signal and
provides equal power to the two antennas.
Antennas.
The two telemetry antennas radiate the telemetry carrier to ground
receiving stations. The antennas are mounted 180 deg apart on
metal panels located on the upper portion of the propellant tanks.
The antennas provide an omnidirectional pattern about the launchvehicle roll axis.
Telemetry Calibrator Assembly.
Periodic calibration of the telemetry systems establishes references
for data reduction to increase the accuracy of data analyzation.
Calibration commands and signals for preflight and inflight cali-
bration of telemetry systems GP1 and GFI are generated by the
telemetry system calibrator. Preflight calibration is controlled from
the LCC. Calibration is accomplished by switching the input of
a data channel from the data source to the calibrator output. The
calibrator supplies a precisely controlled voltage with steps at 0,
25, 50, 75, and 100 percent of channel capacity. When the calibra-
tion command is received, the calibrator will normally cycle to
each step and return to zero. During preflight calibration, however,
the calibrator output can be stopped at any desired step for align-
ment of data amplifiers or other signal conditioning devices.
ENVIRONMENTAL CONDITIONING.
INSTRUMENT COMPARTMENT.Two instrument compartments located above fuel containers F-I
and F-2 require environmental conditioning during preflight opera-
tions only. A ground environmental control system (see Section
VII) supplies the required conditioning medium through swing arm
no. 1 quick-disconnect. Air is used as the conditioning medium
from approximately 1 day and 8 hr prior to liftoff when electronic
components are activated until 30 min prior to liquid hydrogen
loading at 4 hr and 45 min prior to liftoff. GN_ then provides
the inert conditioning required for the remainder of preflight prep-
aration. The conditioning medium is supplied to the instrument
compartments (figure 4-67) through service arm IA, precooling
check valve, upper flexible tubing assembly, and a distribution
manifold in each compartment. The medium is exhausted through
a weldment near the bottom of each compartment, through the
lower flexible tubing assembly, precooling check valve, swing arm
ducting, and into the atmosphere. The compartment inlet temper-
ature is controlled by a sensor which is located in the GSE supply
duct and provides input to the S-IB forward ECS control panel
in the LCC. A bleed orifice on compartment F-l and the two
precooling check valves maintain the desired environment during
flight. At the end of S-IB powered flight for the AS-201 and AS-202
missions, pressure in instrument compartment F-2 was approxi-
mately 6 and 5 psi respectively. Compartment requirements during
preflight cooling are listed in figure 4-68.
Section IV S-IB Stage
POWER DIVIDER
256.2 MHz
PCM/FM
RF COUPLER
TM CALIBRATOR
POWER SUPPLY
240.2 MHz
FM/FM
DDAS
OUTPUT
TO
GROUND
STATION
DIGITAL
DATA
LOW FREQUENCY
ANALOG DATA
LOW FREQUENCY
ANALOG DATA
Y
DATA INPUTS FROM
MEASURI NG SYSTEM
Figure 4-66
HiGH FREQUENCY
ANALOG DATA j
CH.14273-2
4-57
Section IV S-IB Stage
ENGINE COMPARTMENT.
Tall Unit Conditioning and Water Quench.
Four separate manifolds (figure 4-69) provide for distribution of
conditioned air, GNz, GN z deluge, or water quench to the engine
compartments during preflight operations. During normal opera-
tion, the ground environmental control system provides air or GN 2
for compartment conditioning. Air is used as the conditioningmedium until 30 min before lox loading. GN 2 is used for inert
conditioning during the remainder of countdown. In the event of
engine malfunction cutoff or fire in the engine compartment area,
a cold GN 2 deluge is activated flooding the compartments with
cold GN 2. The water quench is available as a backup in case of
fire. Compartment conditioning requirements are presented in
figure 4-70. The conditioning medium, GNz deluge, or water
quench flows through a quick-disconnect coupling, pneumatic
valve, and dispersal manifolds which direct the flow to the enginecompartments and the area between the firewall and the center
lox tank sump. Compartment inlet temperature is controlled by
thermal probes that are located in the engine areas and providetemperature input to the S-IB aft ECS control panel in the LCC.
A pneumatic valve in each manifold is held open during system
operation by 750 psig GN 2 control pressure routed through shortcable mast no. 4 from the launch complex (see Section VII) and
closes at liftoff to prevent exhaust gases from entering the enginecompartments during flight.
Fire Detection.
The fire detection system consists of 32 thermoeouple sensors (con-
nected in four loops of eight sensors) located on the aft thrust
BLEED ORIFICETO ATMOSPHERE "_%.
INSTRUMENT
COMPARTMENT
NO. I
FUEL TANK
F1
MANIFOLD
OUTLET (LOWER)FLEX DUCT ASSEMBLY
LEGEND
CONDITIONED MEDIUM SUPPLY
EXHAUST MEDIUM
VENT
ELECTRICAL
THERMAL
PROI
UMBILICAL
PRECOOLING CHECK
VALVE (2)
LOX TANK
L2
SERVICE ARM
IA
EXHAUSTED
TO ATMOSPHERE
FLEX DUCT ASSEMBLY
/_-.
I NSTRUMENT
COMPARTMENT
NO. 2
TEMPERATURE FEEDBACK TO GSE
FUEL TANK
F2
I
4-58
Figure 4-67
CONDITION MEDIUM (AIR OR GN2)=lll_)
J
CH-14282-1
Section IV S-IB Stage
PROBE DATACOMPARTMENT INLET CONDITIONS COMPARTMENT
FLOW & TEMP. RANGE SETTING
(Lbs/Min) PRESSURE TEMP. HUMIDITY (Degrees F) LOCATION (Degrees F)(In. of H20) (Degrees F) (Gr/Lb of Air)
45 15 75 + 5 0 - 37 #1, 32-140 DUCT 75 + 3
12, 68- 84 (GSE)
NOTE: REQUIREMENTS FOR AIR AND GN 2 ARE SAME.CH.14284-I
Figure 4-68
structure and the firewall substructure. The sensors monitor tem-
perature rise-rates in critical engine areas prior to liftoff. The sensor
output is displayed on a recorder in the LCC where it is continu-
ously monitored by a redline observer. If the display indicates
presence of a fire (a temperature rise of 60 ° F/sec for 0.5 see),the observer notifies the test conductor who initiates a launch scrub
and orders the cold GN 2 deluge turned on. A water quench may
be manually initiated if necessary to extinguish fire.
HAZARDOUS GAS DETECTION SYSTEM.
To detect the presence of hydrogen and oxygen gases, the atmo-
sphere in the S-IB stage engine compartment, the S-IB lox tankskirts, the S-IVB forward skirt, and the S-IVB aft interstage is
monitored during prelaunch operations. This is accomplished
through a hazardous gas analyzer located in the ML that draws
gas samples from the various compartments through a mass ana-
lyzer. See figure 4-71.
ORDNANCE.
Ordnance components used on the Saturn IB vehicle initiate various
operations necessary for proper stage function. Additionally, ord-
nance installed on the S-IB stage propellant containers provides
a propellant dispersion capability for use if the vehicle becomes
a safety hazard during the boost phase of flight.
H-1 ENGINE ORDNANCE.
Four types of ordnance devices either initiate or shut down tlae
H-I engine operation. Those items are the solid propellant gas
generators (SPGG), the SPGG initiators, the squibless igniters, andthe conax valves. See H-I Engine for details on these items.
S-IB STAGE PROPELLANT DISPERSION
SYSTEM.
The S-IB Stage Propellant Dispersion System (PDS) will sever each
of the nine propellant tanks and disperse the propellants if flight
termination becomes necessary. Two exploding bridgewire (EBW)
firing units, a safety and arming device (S&A), two EBW detonators,Primacord, and flexible linear-shaped charges (FLSC) make up
the PDS (figure 4-72). Primacord assemblies interconnect the FLSCto detonators in the S&A device. The EBW firing units interface
with the PDS detonators and the secure range safety command
system. If flight termination becomes necessary, the range safetycommand system (Section I) will provide signals to arm (charge
the EBW high voltage storage capacitor) and trigger the firing units,which deliver high-energy electrical pulses to the EBW detonators
in the S&A device (figure 4-73). Explosive leads in the S&A device
rotor propogate the detonator explosion to the Primacord and
subsequently to the FLSC assemblies. The FLSC assemblies rupture
the propellant tanks allowing the propellants to disperse radially
from the stage and burn rather than explode. The burning propel-
lants result in only a fractional amount of the theoretical yield
if the vehicle should explode. The reliability of Saturn propellant
dispersion systems was demonstrated during the flights of SA-2and SA-3. After S-IB stage engine cutoff a destruct command
destroyed the vehicle to release water ballast contained in the
dummy upper stage (Project Highwater).
EBW Firing Unit.
Exploding bridgewire firing units, used extensively in the Saturn
IB launch vehicle, provide electrical pulses to ignite the ordnance.Redundant EBW firing units in each system ensure the desired
function. The solid-state electronic EBW firing units generate a
high voltage, high energy, short duration pulse to fire EBW detona-
tors or initiators. Each firing unit consists of an input filter, a
charging circuit, a storage unit, a trigger circuit, and an electronic
switch. Two inputs are necessary to the firing unit function. The
first input, the application of 28 Vdc, passes through the inputfilter, which functions as a noise suppressor, to the charging circuit.
In the charging circuit, an oscillator, a step-up transformer, anda silicon controlled rectifier increase the voltage to 2300 Vdc and
couple the charge to the capacitor storage unit. A voltage divider
network provides a 0- to 4.6-Vdc input to the telemetry system
proportional to the 0- to 2300-Vdc charge on the storage unit. The
second input applies 28 Vdc to the trigger circuit. The trigger
disables the charging circuit and generates a signal voltage for the
electronic switch. This permits the storage unit to discharge throughthe switch and the detonator or initiator. A minimum of 1.5 sec
must elapse between inputs.
The S-IB propellant dispersion panel provide control and monitor-
ing of the firing units during checkout and prelaunch operations.
Power switches control power application for ground power opera-tions and internal power operations. POWER ON and ON INTER-
NAL POWER indicators monitor the condition of the firing units.
A FIRING UNIT VOLTAGE meter provides a readout of the
firing unit charge voltage during checkout. A 0- to 4.6-Vdc signal,
proportional to the 0- to 2300-Vdc charge, drives the meters. During
flight the charge monitor signal will be telemetered to ground
receiving stations and recorded.
During checkout operations, the output pulses from the firing unitsare checked by connecting the output cables to pulse sensors. The
firing units are then charged and triggered by the secure range
safety command system. The pulse sensors receive the firing unit
output and provide a 28-Vdc output to the FIRED indicators on
the S-IB propellant dispersion panel. After checkout, the pulsesensors are removed from the vehicle and the firing unit outputcables are attached to the EBW detonators in the S&A device.
See figure 4-74 for firing unit characteristics.
EBW Detonator.
EBW detonators provide the electrical-to-ordnance interface for
explosive fuse assemblies. The detonators function on the EBW
principle: when a high voltage, high energy pulse is applied toa small diameter, low resistance wire element, the element explodes,
rapidly releasing a large amount of energy adequate to detonate
4-59
SectionIVS-IBStage
ENGINECOMPARTMENTCONDITIONING
FLANGE
BARREL
SPERSAL
MANIFOLD
LEGEND
WATER QUENCHSUPPLY
-80°/: GN 2
CONDITIONED
MEDIUM SUPPLY
/ 750 PSIG GN 2
ELECTRICAL
RMAL
PROBES
LOWER
SHROUD
PNEUMATIC
VALVE
Note
THE ASSEMBLY SHOWN IS
TYPICAL IN EACH QUADRANT VEHICLE
GSE
WATER QUENCH SUPPLY
(MANUALLY-INITIATED •BACKUP FOR GN 2
LAUNCHER
MANIFOLD
QUICK-DISCONNECTCOUPLING
TEMPERATURE
FEEDBACK TO
GSE
Figure 4-69
an explosive charge. The detonator consists of an acceleration
charge, a main charge, a bridgewire; two electrical connector pins,and insulators-all enclosed in a steel case. The case has threads
on each end that permit installation into the safety and arming
device and provide for attachment of the EBW firing unit cable
to the detonator. The detonator is not sensitive to static discharges
or RF energy, and the explosives are relatively insensitive to heat.
A spark gap in one electrical connector pin precludes accidental
dudding of the detonator through inadvertent application of ground
power. The gap, created by a thin mica spacer that separates thetwo metal sections of the pin, has a breakdown voltage of 600
to 1200 Vdc. The spacer is square and creates a gap in the four
areas where it does not cover the face of the cylindrical pin sections.
UMBILICAL
PLATE
CABLE
MAST #4
_-80°F GN 2
_ _FROM GN 2 DELUGE SYSTEM
_(IN CASE OF FIRE ONLY)
750 PSIG GN2CONTROL PRESS.
FROM GSE,
CONDITIONED
MEDIUM
AIR OR GN 2
CH-14315-2
4-60
Alumina sleeves and insulators secure the electrical connector pins
in the detonator. The bridgewire connects between electrical pins
at the base of the acceleration charge. The main charge is located
in the end of the detonator just forward of the acceleration charge.
The main charge is located in the end of the detonator just forward
of the acceleration charge. Closure disc paper separates the charges
until detonation. The 2300-Vdc pulse from the EBW firing unit
melts the fine bridgewire. The magnetic field created by the currentcauses the melting wire to form a series of lobes. When the lobes
become unconnected spheres of molten metal, arcing occurs be-
tween the spheres. Very high internal pressure within the molten
spheres explosively propels hot particles into the 1.04-+-0.03 gr,
Class 2 PETN, acceleration charge. The energy release and shock
!NGINE COMPARTMENT CONDITIONING REQUIREMENT
FLOW
MEDIUM
AIR
GN 2
GN 2 DELUGE
WATER QUENCH
FLOW
(llx/mln)
135--170
30O
42O
8000 g_n
COMPARTMENT INLET CONDITIONS
I_ESSURE
(In of 1"120)
42
42
84
125 I_lg (SUPPLY)
TEMP
(DEGREES E)
40-150
65-95
-80
AMBIENT
HUMIDITY
(Gt/Ib of a_r)
O-37
O-37
O-3"/
N/A
N/A - NOT APPLICABLE
Section IV S-IB Stage
COMPARTMENT
TEMP RANGE
(DEGREES F)
50"__10
40-75
N/A
N/A
PROBE DATA
SETTING
LOCATION (DEGREES F)
STAGE 50__10
STAGE 50"__10
STAGE N/A
STAGE N/A
CH.14283-1
Figure 4-70
of the exploding bridgewire is sufficient to initiate PETN detona- that arms the PDS, fulfills the requirements stated in AFETRM
tion. The acceleration charge then detonates the main charge con- 127-1 for arming a flight termination system. The specific require-
sisting of 1.40_0.25 gr of Class 4 PETN. ments are:
Safety and Arming Device, a. The unit must complete and interrupt the explosive train by
The safety and arming (S&A) device, an electro-mechanical device remote control.
WATER=_o_
IIIIL
1
I
SERVICE ARM 6
IU
S-IVB FWD SKIRT
S-IVB STAGE
AFT INTERSTAGE
Figure 4-71
4-61
Section IV S-IB Stage
S-IB STAGEPROPELLANTDISPERSIONSYSTEM
LOX PRESSURIZATION
TYP
1--
3O°TYP
SPIDER
AERODYNAMIC
FUEL
PRESSURIZATION
MANIFOLD
VIEW LOOKING AFT
FUSE ASSY
(4 PLACES)
VALVE
CHARGE
(9 PLACES)
VENT
VALVE
AERODYNAMIC
EBW
DETONATORS
FIRING UNIT
EBW
FIRING
N. _-_. LOX TANK
NO, 3" '---_ TYPICAL
SHAPED
CHARGE
AND
ARMING DEVICEi
FUSE ASSY
PRIMACORD
BLOCK_
29" CONDUIT SECTION
(22 SECTIONS ON EACH
OUTER TANK, 9 SECTIONS
ON CENTER TANK)
\
FIN LINE 4
/
PRIMACORD
LEADS
TANK STATIONS) ,.,,
_'_ _"-'-IWO LEADS TOSHROUDFUSEASSY
-- _,,_-60--GPF PRIMACORD
_ j_COUPLINGl"- CLAMP
_ 100-GPF SHAPED,._ CHARGE
_ F SPLICE BLOCK
SEALANT COATING
BETWEEN SPLICE BLOCKAND INSULATION
L'-" _--'=_T-""J_ INSULATION
DETAI L B
r-- LOCKWIRE
RUBBER CONDUIT"_ /-- SHIELD WELDMENT
INSULATION._ \J, FBOLT
SHAPED CHARGE _TANK SKIN
_----TWO LAYERS OF
FILAMENT TAPE
SECTION A-A
Nete
:ONDUIT AND INSULATION
PROTECTS THE SHAPED CHARGE
FROM AERODYNAMIC LOADINC
AND HEATING. INSULATION
ALSO MAINTAINS THE SHAPED
CHARGE AT PROPER STANDOFF
DISTANCE FROM TANK SKIN TO
MAXIMIZE THE CUTTING EFFECT
SECTION
tA-A
STATI ON
288.901
(SEE TABLE
FOR OTHER
TANK STATIONS)
S-IB STAGE SHAPED CHARGE STATION LOCATIONS
TOP OF ASSY. BOTTOM OF
CONTAINER TOP OF (SPLICE BLOCK ASSEMBLY
CONTAI NER TOP) (ADAPTER END)
105" LOX 941.304 939.119 693.524
LOX NO. 1 937.961 935.961 288.901
LOX NO. 2 937.961 935.961 288.901
LOX NO, 3 937.961 935.961 288.901
LOX NO. 4 937.961 935.961 288.901
FUEL NO. I 935,429 933.369 283. 619
FUEL NO, 2 935.429 933.369 283. 619
FUEL NO. 3 935,429 933.369 285.369
FUEL NO. 4 935.429 933.369 285.369
CH-14203-2
4-62 Figure 4-72
Section IV S-IB Stage
b. The unit must provide indications to remote monitoring equip-
ment whether its position is safe or armed.
c. Visual position indication and manual operation of the unit must
be possible.
d. Must have separate connections for the arming and firing cir-cuits.
The S&A device consists basically ofa Ledex 95-deg rotary solenoid
assembly, a metal rotor shaft with two 2.5-gr PETN explosive
inserts, and position-sensing and command switches that operatefrom the rotor shaft cam. On electrical command from the ground
system just prior to automatic countdown, the solenoid assemblyrotates the shaft containing the two explosive inserts 90 deg. This
aligns the inserts between the EBW detonators and the Primacord
adapters to form the initial part of the explosive train.
Each power application to the solenoid assembly moves the rotor
90 deg in one direction, thus rotating the explosive inserts in and
out of alignment with the explosive train. A spring-loaded detent
locks the rotor shaft in position; a spring and ratchet arrangement
returns the solenoid to the original position. When the S&A device
EBW FIRING UNIT CHARACTERISTICS
PROPELLANT DISPERSION
SYSTEM DIAGRAM
ROTOR
SAFE
END VIEW
EBW FIRING
UNIT CABLES _
EBWEXPLOSIVE I_ I1_ F DETONATORLEAD (2)
AFETY AND
ARMED
IV ARMING
s_ A DEVICE/11 DEVICE
F2 L///_100 GPF SHAPED
CHARGE (9 PLACES) -..-/ C.H 14202-1
SAFE ANI ARM DEVICE CHARACTERISTICS
Figure 4-73
PARAMETER
INPUT POWER
pEAK CURRENT INPUT
STEADY STATE AVERAGE
CURRENT INPUT
OUTPUT VOLTAGE
CHARGE TIME
TRIGGER INPUT VOLTAGE
TRIGGER CKT SENSITIVITY
TRIGGER CKT RESPONSE
TRIGGER PEAK CURRENT
TRIGGER STEADY STATE
AVERAGE CURRENT
OPERATING TEMP
COMPONENT LIFE
VALUE
24 TO 32 Vclc
2.0A
250 mA MAX
2,3(]0+ 100 Vdc
1.5 see MAX
24 TO 32 Vdc
a. 8 Vdc OR LESS WILL NOT
ACTUATE TRIGGER
CIRCUITRY
b. A 50 Vdc TRANSIENT PULSE
OF 50 _SEC DURATION
WILL NOT ACTUATE THE
TRIGGER CIRCUITRY
4+1 mS
250 mA
75 mA MAX
-65°F TO +200°F
CONTINUOUSLY 'ON, * THE
FIRING UNIT WILL PERFORM
1000 OPERATIONS, RECEIVINGA TRIGGER AT 15 MIN INTERVALS.
Figure 4-74
CH-14427.2
is in the safe position, the explosive inserts are 90 deg out of
alignment with the EBW detonators and Primacord adapters; therotor forms a metal barrier to prevent propagation to the explosive
train should the detonators inadvertently fire. As an additional
safety precaution to prevent damage to the explosive train, ventholes are drilled through the S&A device body, in line with the
safe position of the explosive inserts, to vent detonation gasesoverboard. A debris retainer instzlled over the vent holes will
prevent damage to surrounding equipment should the detonatorsfire. The vent holes have no function when the S&A device is
ITEM
EXPLOSIVE INSERTS
OPERATIONAL TEMP
RESPONSE TIME
IMPACT SENSITIVITY
DESCRIPTION
TWO 2.5 _ PETN LEADS-65OF TO +165°/:
ROTOR SHAFT WILL ROTATE
90 DEG UNDER NO LOAD IN
LESS THAN 50 MSEC WITH
28 _+4 Vd¢ APPLIED TO SOL
THE S&A DEVICE HAS BEEN
QUALIFIED THROUGH 8-FOOT
DROP TESTS AND SHOCK
TESTS OF 35 g MAGNITUDE,
THREE SHOCKS IN EACH OF
THREE AXES. SHOCK WAVES
WERE I/2 SINE WAVES, 8 MSECDURATION
Figure 4-75
C-H 14428-1
4-63
Section IV S-IB Stage
armed. The S&A device electrical circuitry provides monitoring
and actuation functions for the device, but contains no provisionsfor initiating the explosive train. The position of the device safe
or armed can be monitored remotely (by electrical signals frominternal sensing switches'prior to liftoffonly) on the S-IB and S-IVB
propellant dispersion panel indicators. A safe/arm switch on the
panels permits manual control of the S&A device position for
checkout and arms the S&A device approximately 4 min before
liftoff. A clear polystyrene plug in the end of the S&A device (at
the end of the rotor shaft) provides a means of locally monitoring
the shaft position. The plug is removable, providing access to
manually reset the S&A device to the safe position, if necessary.
See figure 4-75 for S&A device characteristics.
PDS Ordnance.
In addition to the EBW firing units, EBW detonators, and S&A
device, the S-IB stage PDS has an adapter fuse assembly, a shroud
fuse assembly, and nine FLSC assemblies. The adapter fuse assem-
bly consists of two parallel 6 l-in. lengths of60-gpf PETN Primacord
with end fittings adhesively bonded to the Primacord on one end
of the assembly. The end fittings connect the fuse assembly to
the S&A device and contain 6-gr PETN booster charges to ensure
propagation of the detonation across the mechanical connection
to the Primac0rd. The adapter fuse assembly extends to the shroud
fuse assembly and connects to the shroud fuse assembly by an
overlapping splice. In this connection, the ends of the Primacord
are placed parallel to the shroud fuse assembly, and both are taped
together. See figure 4-76 for PETN characteristics.
The shroud fuse assembly installed on the aft interstage aerody-
namic fairing encircles the forward end of the S-IB stage, serving
as the main explosive train to which the nine FLSC assemblies
connect. The fuse assembly is a 835-in., 60-gpf length of Primacord
wrapped with aluminum tape. Quick-release clamps secure the fuse
assembly to the aft interstage aerodynamic fairing extension.
An FLSC assembly consisting of a lead sheathed 100-gpf PETN
shaped charge, a splice block, and three pieces of 60-gpf Primacordis mounted on each propellant tank. The FLSC assemblies extend
the entire length of the four fuel and four outer lox tanks. Thecenter lox tank FLSC assembly extends approximately half waydown the tank. The FLSC assemblies are inserted into conduits
attached to the tank skin. The conduits provide protection against
aerodynamic loading and maintain the correct orientation of the
shaped charge to the tank skin. The splice block on each FLSC
assembly attaches the Primacord leads to the FLSC. Inside the
splice block, one end of the 5-in. Primacord length butts against
the end of the FLSC. The two longer Primacord lengths, that
connect the FLSC assembly to the shroud fuse assembly, overlap
the short Primacord to FLSC splice. An adhesive, which fills the
cavity around the Primacord and FLSC in the splice block, and
a cover fastened to the splice block secure the Primacord-to-FLSC
connection. The parallel Primacord leads extending from the FLSC
assemblies to the shroud fuse assembly are 30 in. long for the
fuel tanks, 50 in. long for the outer lox tanks, and 150 in. longfor the center lox tank. All the Primacord leads attach to the shroud
fuse assembly by overlapping splices.
POWER TRANSFER SAFETY SWITCHES.
Two squib-actuated switch assemblies, installed on Power Distrib-
utor 12A25 in unit 12 of the S-tB stage, actuate at liftoff to assure
application of internal power to stage electrical systems during
flight. The one-shot switches parallel the contacts of the command
power transfer relays and provide relay contact chatter compensa-
tion by maintaining power application to the energizing coils of
the power transfer internal relays. Each switch assembly consists
of an electrical connector, an adapter, wiring, and a squib-switch.
Potting compound encapsulates the squib-switch and wiring leaving
only the switch terminals exposed for test purposes. A plastic tube
PETN CHARACTERISTICS
SAFETY SWITCHES CHARACTERISTICS
4-64
DESCRIPTION
CRYSTAL DENSITY
MOLECULAR WEIGHT
CHEMICAL FORMULA
MELTING POINT
DETONATION RATE
HEAT OF COMBUSTION
FIRE HAZARD
IMPACT SENSITIVITY
ICC CLASSIFICATION
SYNONYM
SPECIFICATION
WHITE CRYSTALS
1.765
316.55
C(CH2NO3)4
280°F TO 28501:, DECOMPOSES RAPIDLYABOVE 41001:
27,232 ET/SEC AT 1.70 GRAM/CC (I-INCHDIA SAMPLE)
1974 CALORIES/GRAM
MODERATE, BY SPONTANEOUS CHEMICALREACTION
4.4 LBM BUREAU OF MINES (20 MG SAMPLE)
6.6 IN. PICATINNY ARSENAL (16 MG SAMPLE)6 INCHES.
CLASS A
PENTAERYTHRITOL TETRANITRATE
MIL-P-387
C-H 14476-1
Figure 4-76
encloses the assembly to protect the exposed terminals. The switch
assembly has three normally open contacts and one normally closed
contact. Firing current is applied to the squib bridgewire through
the normally closed contact, which opens when the switch actuates.
This action removes power from the squib circuit to prevent possible
power drain after the switch actuates. Switch actuation is accom-
plished by gas pressure generated by the squib. The ICC does
not require special handling of the switch assembly because of
the small squib charge (approximately 100 mg). The squib has
been used extensively in aerospace programs without any known
ruptures of the case. The squibs have been qualified per Picatinny
Arsenal X PA PD 2145 and per Naval Ordnance Lab. OS 10077,
OS 10076, NOTS XS417. See figure 4-77 for safety switches charac-teristics.
ITEM
GAS GENERATING CHARGE
CHEMICAL COMPOUND
QUANTITY
BRIDGEWIRE RESISTANCE
FUNCTIONING TIME
MAX NO-FIRE CURRENT
MIN ALL-FIRE CURRENT
CLOSED CONTACT CAPACITY
OPERATING TEMPERATURE
DROP TEST
DESCRIPTION
LEAD MONONITRORESORCINATE (90%) AND
POTASSIUM CHLORATE (10%)
120 MG. (MAX)
1.8_+ 0.2 OHMS (SWITCH UNASSEMBLED),
1.3 + 0.3 OHMS (SWITCH ASSEMBLED, WITH
SQU_B LEADS SHORTENED).
10 MSEC (MAX)
O.IOA
I.OA
12 A FOR 6 HOURS; 200 A FOR 100 MSEC
(SWITCHES WITHOUT HEAT SINK, AT +160°F
AND ALL CONTACTS IN SERIES).
-65 TO +160o1:
SQUIB SWITCH SUCCESSFULLY COMPLETED
40-FOOT FREE FALL DROP TEST ONTO A STEEL
pLATE (207 BRINELL HARDNESS) HAVING A3-.INCH THICKNESS PER MIL-STD-352.
Figure 4-77
C-H 14472
SECTION
S-IVB STAGE
TABLE OF CONTENTS S-IVBCUTAWAYIntroduction .................................................................... 5-1
Structure .......................................................................... 5-1
Propulsion ...................................................................... 5-8
Pneumatic Control System ............................................ 5.34
Flight Control .................................................................. 5-36Electrical .......................................................................... 5-56
Instrumentation Systems ................................................ 5-61
Environmental Conditioning .......................................... 5-64Ordnance ........................................................................ 5-64
INTRODUCTION.
The Second stage, S-IVB, provides thrust from just after first stage
burnout and separation, which is approximately 2 min 30 sec afterliftoff; thrust is continued until orbital velocity is achieved approxi-
mately l0 min after liftoff. During first stage powered flight, the
S-IVB stage along with the aft interstage (figure 5-1) provides the
load-bearing structure between the S-IB stage and the IU. The
propellant tank assembly, the forward and aft skirts, and the thruststructure form the basic stage structure. One J-2 engine mounted
on the thrust structure propels the vehicle during second stage flight
with 225,000 lbf thrust. The propellant tank assembly is composed
of lox and LH 2 tanks separated by a common bulkhead. Most
of the stage electrical equipment is mounted around the innersurfaces of the forward and aft skirts (figures 5-2 and 5-3). Each
skirt also has an umbilical interface plate for prelaunch ground
connection to the various stage systems. Two auxiliary propulsion
system modules, mounted on the aft skirt exterior, provide pitch,
yaw and roll attitude control during orbital coast periods and rollcontrol during S-IVB burn. The engine-mounted hydraulic system
is used for J-2 engine pitch and yaw gimballing to provide attitude
control during S-IVB powered flight. The S-1B and S-IVB stages
separate at the S-IVB aftskirt and aft interstag¢ joint. The three
ullage rockets and four retromotors, which aid this separation,
mount respectively on the aft skirt and the aft interstage exteriors.
Plumbing and wiring that must pass outside the S-IVB stage are
covered by the main tunnel, the auxiliary tunnel, and various
fairings on the aft skirt. Figure 5-4 lists a summary of basic stage
and stage systems data.
STRUCTURE.
Figure 5-5 shows the primary, load-carrying structural assemblies
of the S-IVB stage consisting of the propellant tank, forward and
aft skirts, engine thrust structure, and aft interstage in addition
to the non-primary structural elements consisting of the main
tunnel, auxiliary tunnel, and aerodynamic fairing segments. The
primary structural purpose of the stage is to transfer loads imposed
by the spacecraft and instrument unit to the S-IB stage as well
as maintaining structural integrity for the additional loads generated
by the S-IVB stage itself. Local shell structure must also have the
ability to support the loads caused by mounting of equipment,
protuberances, and externally located systems. The required ul-
timate safety factor of the stage of 1.4 has been verified by analysis
FWD
LH
FWD SAFETY
ANTENNAS (2)
UMBILICAL
ANTENNAS (4)
.IARY
"fUNNEL
N TUNNELCYLINDRICAL
TANK kJ
He SPHERES
LOX TANK PRESS.(6)COMMON
CHILLDOWN RETURN LLAGE
ROCKET (3)
APS MODULE DOME
LH 2 FEED UMBILICALDUCT
EMENT
AFT SKIRT CURTAIN
LH2 STRUCTURE
RETROMOTOR
_ITERSTAGE
ENGINE
AERODYNAMIC
CH-14478-1
Figure 5-1
and tests. All possible critical shell load conditions including accept-
ance firing, prelaunch, boost, stage powered flight, and orbitalconditions have been investigated. This investigation has established
that the structural shell is, in general, critical for the boost condition
at the time of maximum aq. The design requirement for each
structural assembly, the validity of the design requirement, and
the ability of the structural assembly to meet the design requirementhave been established. The measured weight of the S-IVB-206 stage
(including the S-IVB flight interstage) is 27,771 Ibm and it measures
260 in. in diameter by 700 in. in length.
The S-IVB structural material selections were predicated on high
strength and cryogenic capabilities, advances in welding and bond-
ing techniques, and a history of previous and similar applied me-chanics. The material stock used were casting, sheets, plates, ex-
trusions, forging, and honeycomb. The joining methods were weld-
ing, riveting, lockbolting, bonding, and bolt nut/insert removeable
5-1
I
Section V S-IVB Stage
DETECTOR
DETECTOR
COUPLER
RF ASSY
DIRECTIONAL
HYBRID POWER
DIVIDER A34 "_
PANEL NO.14
41tA
LOAD
A203
/A _ SWITCH A202
NEL NO.AI230lOWER DIVIDER
411A64
ASSY A200
-MOUNTING
BRACKET
A202
LIQUID LEVEL SENSOR
CONTROL UNIT A27-._
LIQUID LEVEL SENSOR
CONTROL UNIT A26"_ \
L,QU,DLEVELSENSOR\ \CONTROLUN,TA2S--_\ \
LIQUID LEVEL SENSOR \ \ \
CONTROLUN,TA2,---_\ \ \_.QUIDLEW,SENSOR\
CONTROL_4
_ INVERTER
] _],_ CONVERTER
_'_'r_ _ ELECTRONIC ASSY A7
"L'_ _ _ PROPELLANT UTILIZATION
ELECTRONICS ASSY A6
_--- PANEL NO.II -41 IA92
SIGNAL CONDITIONING
RACK 41 IA61
(PANEL POSITION NO. IC
COMMAND
RECEIVER NO. 2 A18
SAFETY
COMMAND
RECEIVER
NO.1 AI4
SAFETY
CONTROLLERASSY NO.2 RS EBW FIRING
A19 UNIT
RS EBW FIRING
RANGE UNIT NO. ISAFETY
5VDC EXCITATIONCONTROLLER IU CONNECT
ASSY NO. 1 ELECTRICAL
A 13 RECEPTACLE BUS MO DULE
SUPPORT AI A200
5VDC
EXCITATION
MODULE A2
CENTRAL
DECODER
DECODER BUS A
PANEL NO. 15
v BUS
)RWARD
SKIRT
(VIEW LOOKING
AFT)
28V FWD NO.2
BATTERY 41 IAI 1
PANEL NO.16
DUMMY
RECEPTACLE
RSSAFETY
AND ARMING
- RS DECODER
NO.2 A2
NO. 1 AI
28V FORWARD
POWER DISTRIBUTION
MOUNTING ASSY AI0
THERMO-CONDITIONING SUPPLY
LINE TO INSTR
5-2
Figure 5-2
FWD CONTROL DISTRIBUTION
MOUNTING ASSY 41 IA9
GROUND PLATE
41 IA36
WD NO. 1
BATTERY 411A8 (2 PLACES)
THERMO-CONDITIONING
RETURN LINE FROM
INSTRUMENT UNIT CH.14544.2
Section V S-IVB Stage
AFT SIGNAL
CONDITIONING
RACK 404A62
(PANEL POSITION NO.
AFT SIGNAL CONDITIONING
RACK 404A63
(PANEL POSITION NO.9
LIQUID
LEVEL
CONTROL
UNITS-- PANEL NO.6
404A71
)DER BUS S
MODULE A200
PANEL
404A72
DECODER
A200
7
9
CALSIGNAL
BUS MODULE
A201
ATTITUDE CONTROL
RELAY MODULE NO. 2
A19
AFT SKIRT
(VIEW LOOKING AFT)
RESISTOR
BUS MODULE
SIGNAL GROUND
BUS MODULE
ESISTOR
BUS MODULE
,IA205
BUS MODULE
PANEL NO.4
404A61
NO.3
404A60
CP180
MULTIPLEXER
ASSY A200
MULTIPLEXER
ASSY A201
ECTRICAL
DISTRIBUTION
MODULE BRACKET
A204
REMOTE DIGITAL
SUBMULTIPLEXER
A200
4ALOG
SUBMULTtPLEXER
A201
Figure 5-3 (Sheet 1 of 2)
attachments. Aluminum alloys (2014-0 heat treated to 2014-T6,2014-T651, and 2014-T652) were used in fabricating the welded
propellant tank. The skirts, thrust structure, and aft interstageassemblies are skin-stringer-ring frame fabrication. Aluminum
alloys (7075-0 heat treated to 7075-T6, 7075-T6 and A356) were
used and joined by mechanical fasteners. The common bulkhead
is a honeycomb composite structure constructed of welded spherical
SEQUENCER MOUNTING ASSY
404A3 (PANEL POSITION NO. 1)
SELECTOR 404A1
PANEL NO.2
404A70
CH.14,554.1
formed 2014-T6 aluminum alloy face assemblies bonded to heat
resistant phenolic honeycomb core material.
PROPELLANT TANK STRUCTURE.
The propellant tank structure shown in figure 5-6 consists of a
268 in. long cylindrical section constructed by butt welding longi-
5-3
Section V S-IVB Stage
AFT CONTROL DISTRIBUTION
MOUNTING ASSEMBLY 404A4
(FANEL POSITION NO, 20)
28V POWER DISTRIBUTION
MOUNTING ASSEMBLY 404A2
(PANEL POSITION NO. 19)
28V AFT NO. 1
BATTERY,404A1
CHANNEL DECODER
ASSY.
EI._ECTRICAL
CONNECTOR
BRACKET, A20t
GROUND PLATE,
PANEL NO. 18,
404A75
JETTISON
EBW FIRJNG
UNIT NO. 1,
A1
EBW FIRING
UNIT NO. 2,
A2
)NNECTOR
MODULE, A200
CONTROL
RELAY MODULE,
1, A4
5V EXCITATION
MODULE, A7
UR JETTISONDETONATOR
BLOCK, A3
AFT SKIRT
(VIEW LOOKING
AFT SIGNAL CONDITIONING
RACK, 404A64
(PANEL POSITION NO.
GROUND PLATE,
404A57
PANEL NO. 15
5-4
56V AFT NO. 2
BATTERY, 404AI I
56V AFT
POWER DISTRIBUTION
MOUNTING ASSY.
404A45
(PANEL POSITION 17)
LOX CHILLDOWN
INVERTER ASSY A1
Figure 5-3 (Sheet 2 of 2)
FUEL CHILLDOWN
INVERTER ASSY A2
CH-14543-1
Section V S-IVB Stage
fA SUMMARY;-IVB ;TAG
DIMENSIONS
LENGTH
DIAMETER
MASS
DRY STAGE
LOADED STAGE
AT ORBITAL INJECTION
. S-IB/S-IVB INTERSTAGE
59.1 FT
21.7 FT
22,150 LB m
256, 800 LBm
28,200 LBm
5,661 LBm
LLAGE ROCKETS
NUMBER OF ENGINES
THRUST (PER ENGINE)
EURNTIME
PROPELLANT
LOCATION
PRESSURIZATION SYSTEM
3
3,460 LBf@ 1,000,000 PT
3.9 5EC
SOLID
120 DEG INTERVALS AROUND S-IV8 AFT SKIRT.
ENGINES CANTED OUTWARD 35 DEG.
JETTISONED 15 SEC AFTER STAGE SEPARATION.
PROPELLANT LOAD
ROCKET ENGINES
J-2 ENGINES
BURNTIME
THRUST
PROPELLANT
MIXTURE RATIO
EXPANSION RATIO
232,200 LB m
440 SEC
225,000 LBf
@ 200,000 FT
LOX AND LH 2
5.5:.1 (MAX), 4.8:1 (MIN)
27:1
OXIDIZER CONTAINER
FUEL CONTAINER
OXIDIZER PRESSURE
PREFLIGHT
]NFLIGHT
FUEL PRESSURE
PREFLIGHT
INFLIGHT
HELIUM
INITIAL: HELIUM FROM GROUND SOURCE, GH 2FROM J-2 ENGINE DURING S-IVB BURN
37 TO 40 psia
37 TO 40 psi=
31 TO 34 psia
26.5 TO 29.5 psla
OXIDIZER NPSH
FUEL NPSH
GAS TURBINE PROPELLANT
FUEL TURBINE SPEED
OXIDIZER TURBINE SPEED
HYDRAULIC SYSTEM
ACTUATORS
GIMBAL ANGLE
GIMBAL RATE
35 PSIA MIN
27 PSIA MIN
LOX AND LH 2
27,000 RPM
3,600 RPM
HYDRAULIC (TWO PER ENGINE)
+ 7 DEG SQUARE PATTERN
8 DEG,P3EC EACH PLANE
ENVIRONMENTAL CONTROL SYSTEM
PREFLIGHT AIR CONDITIONING
PREFLIGHT GN 2 PURGE
FLIGHT
ASTRIONICS SYSTEMS
GUIDANCE
TELEMETRy LINK
ELECTRICAL
AFT COMPARTMENT AND FORWARD SKIRT
AFT COMPARTMENT AND FORWARD SKIRT
UNIT CONDITIONING SYSTEM
PATH ADAPTIVE GUIDANCE MODE THRU THE
IU DURING S-IVB BURN
258.5 MHz
BATTERIES: 28 Vdc (3 ZINC SILVER-OXIDE)
56 Vd¢ (I ZINC SILVER-OXIDE)
GIMBAL ACCELERATION
APS
NUMBER OF ENGINES
THRUST (PER ENGINE)
PROPELLANTS
PROPELLANT LOAD
171.5 DEG,ASEC 2
6 (3 PER MODULE)
150 LRf VACUUM
HYPERGOLIC (MMH AND N20 4)
62 LBm PER MODULE
STATIC INVERTER-CONVERTER: 28 Vd¢ TO
115 Va¢, 400 Hz SINGLE PHASE, 25 Vpp,
400 Hz SQUARE WAVE AND 117, 21, AND
44.2Vd¢.
CHILLDOWN INVERTER (2): 56 Vd¢ TO 56 Vac,
3 PHASE, 400 Hz QUASI-SQUARE WAVE
5-VOLT EXCITATION MODULE: 28 Vdc
TO 5 Vdc, -20 Vd¢ AND 10 Vpp, 2000 Hz
SQUARE WAVE
20-VOLT EXCITATION MODULE: 28 Vd¢ TO
20 Vck:
LOCATION
RETROMOl_ORS
NUMBER OF ENGINES
THRUST (PER ENGINE)
BURNTIME
I_ROPELLANT
LOCATION
AFT SKIRT AT FIN POSITIONS I AND III
4
36,720 L8f@ 200,000 FT
1.52 SEC
SOLID, TP-E-8035 (THIOKOL)
90 DEG INTERVALS
RANGE SAFETY SYSTEM PARALLEL ELECTRONICS, REDUNDANT
ORDNANCE
Note
ALL MASSES ARE APPROXIMATE
Figure 5-4
CH-144S3-2
tudinally the seven cylindrical section segments. The cylindrical
section segments are made from 0.750-in. plates, which are first
mechancially milled to make the waffle pattern, and then brake-
formed to a 130-in. radius. The waffle pattern consists of pockets
9.5 in. on center that are oriented ___45 deg with respect to the
vehicle longitudinal axis. These pockets are surrounded by ribs
that are 0.627 in. high and 0.144 in. wide. The skin or web thickness
at the bottom of the pocket is 0.123 in., and the weld land areas
at the edges of these segments are 0.252 in. thick. The forward
dome is fabricated from nine "pie-shaped" gore segments welded
together and to the jamb manhole, which provides access to the
LH 2 tank. These segments are made of 0.150-in. thick aluminum
alloy sheet stock, which is formed to a 130-in. radius then chemicallymilled to a skin thickness of 0.060 in., with a thickness building
5-5
Section V S-IVB Stage
up to 0.118 in. at the weld-land areas. The increased weld-land
thickness is for the purpose of maintaining the stress levels below
the yield point of the material in the "as welded" state.
The aft dome, like the forward dome, is fabricated from nine
"pie-shaped" gore segments. These segments are made from 0.280-
in. thick sheet stock and are chemically milled after forming. Theweld-land areas, including the rings around the dome where the
common bulkhead and the thrust structure attach, are milled to
0.191 in. The liquid hydrogen portion of the dome (the area forward
of the common bulkhead joint) is chemically milled into a waffle
pattern with a web thickness of 0.082 in. The rib height is a
minimum of 0.245 in. The 0.625-in. wide ribs are located approxi-mately 5.5 in. on center at the maximum, and are oriented at 0
deg and 90 deg with respect to the vehicle longitudinal axis. The
S-IVBSTRUCTURALASSEMBLIES
FORWARDcS2--- J,gtlll!tllllMl#/JOtAOFEN_NGI_
AUXILIARY
TUNNEL
(NON-PRIMARY _._STRUCTURE)
\PROP
TANK
MAI N
TUNNEL
(NON-PRIMARY
STRUCTURE)
STRUCTURE
BEAM (8)
AERODYNAMIC
FAIRING
SEGMENT (8)
(NON-PRIMARYSTRUCTURE
I NTERSTAGE
5-6
Figure 5-5
web area is milled to a thickness of 0.086 and 0.092 in. for the
areas forward and aft of the thrust structure joint, respectively.
Access to the liquid oxygen tank is through the sump jamb inthe bottom of the dome.
Many honeycomb composite structures are used in the vehicle,
the most unique is the common bulkhead that provides both the
structural and thermal separation between the LH 2 and the loxtanks. The common bulkhead forward and aft face sheets are similar
in construction to the other domes, i.e., from nine chemically milled,
welded pie segments. The circumferential edges of the faces are
butt-welded to "Y-shaped" extruded rings, which attach the bulk-
head to the aft dome. The two face sheets are separated by and
adhesively bonded to reinforced plastic honeycomb core material.
The extruded Y-rings are joined together by a common seal weld
and are attached to the aft dome by mechanical fasteners and
by the LH 2 tank and lox tank seal welds.
As stated, all propellant weld areas are thicker than the basic
membranes to appropriately reduce stresses transferred throughthe welds and to allow for lower material allowables due to the
welding process. All welds are inspected by use of dye penetrant
and X-ray techniques. The final quality control operation on the
propellant tank consists of a hydrostatic proof test after which both
weld inspection and tank leak checks are performed. One significant
problem involving the propellant tank jamb welds was encountered.
After hydrostatic proof testing of the S-IVB-502 tank, a few small
cracks were discovered in the jamb weld of the lox tank. It was
established that these areas had been repaired during manufacture.
As a result of this conclusion, all vehicles already manufactured
were inspected in the jamb areas of both the lox and LH _ tankdomes. Small cracks were found in S-IVB-201, -202, and the bulk-
head test specimen. It was decided, even though all stages had
not developed cracks and had successfully passed hydrostatic proof
testing, that it would be advisable to reinforce the jamb weld areas
on all stages. This was accomplished in two phases. The rein-
forcement of the earlier stages including S-IVB-206 was accom-
plished through the addition of internal and external doublers over
the jamb weld area of both the lox and LH 2 tank domes. Thedoublers were attached with mechanical fasteners and also adhe-
sively bonded. This jamb weld reinforcement is shown in a detail
view in figure 5-6. Two 5.5-ft domes, including a jamb weld and
its reinforcement, were made to the lox dome and to the LH z
dome configurations. The two domes were then tested to planned
destruction to determine their structural capability. Both analysis
and test results established that the structural capability was well
in excess of the design requirements. In addition, leakage checks
were performed to provide assurance that manufacturing capability
was adequate to provide a leak-free assembly. The second phase
of the reinforcement program consisted of redesign of the domesand utilizing thicker weld lands to eliminate the need for the
doublers. The redesign was first accomplished on the S-IVB-209LH 2 dome and the S-IVB-210 lox dome.
FORWARD SKIRT, AFT SKIRT, AND AFT INTER-STAGE.
These three assemblies, similar in configuration, are cylinders con-
structed of 7075-T6 aluminum alloy skin, external stringers, and
internal frames. Their diameter is 260 in., but their lengths vary.The forward skirt has 108 stringers while the aft skirt and aft
interstage each have 112 stringers. The forward skirt assembly
extends forward 122 in. from the forward skirt attaching ring ofthe propellant tank structure. The forward skirt provides an attach-
ment plane for the instrument unit and mounting provisions for
electronic equipment on thermo-conditioning panels. After vehicle
assembly on the launch pad, the instrument unit access door pro-
vides entry to the forward skirt interior. The aft skirt assembly
extends aft 85.5 in. from the aft skirt attaching ring of the propellant
Section V S-IVB Stage
JAMB WELD REINFORCEMENT
MECHANICAL
FAS
\ .................. _ JAMB
_'--_ GOR E SEGMENT
FORWARD DOME GORE SEGMENT DETAIL
FOWARD METAL
FACE SHEET (9)
LH 2 TANK
PoLSE::TW2 7
SEALING
COMPOUND _
H_K BOLTJ
AFT DOME /
STRUCTURE----*
CYLINDRICAL
SECTION
SEGMENT (7)_
LH 2 TANK
HEAT
RESISTANT
PHENOLIC
HONEYCOMB COREAFT DOME
GORE SEGMENT (9) --j
AFT METAL
• PACE SHEET (9)
_LOX rANK
SEAL WELD
COMMON BULKHEAD DETAIL
WAFFLEPATTERN
BULKHEAD
{I-H2 - LOX)
ATTACHING RING
ATTACHING RING
LOX SUMP
ELD
INTERNAL DOUBLER
(TYPICAL BOTH DOMES)
K
Figure 5-6
CH-12313-2
5-7
Section V S-iVB Stage
tank structure to the aft interstage. Ullage rockets and the auxiliary
propulsion system mount on the aft skirt exterior; electrical equip-
ment panels mount on the skirt interior. The aft interstage assembly
is 224.5 in. in length and provides the interface between the S-IVB
stage and the S-IB stage. The aft interstage has eight verticalinternal reaction beams that translate the loads to the S-IB 220-in.
diameter structure points. Four retromotors are mounted externally.
The separation plane for the S-IVB stage is the aft end of theaft skirt.
S-IVB THRUST STRUCTURE ASSEMBLY
THRUST STRUCTURE.
Figure 5-7 shows the truncated-cone-shaped thrust structure. The
configuration of the thrust structure is similar to the skirt assemblies
in that 7075-T6 aluminum alloy skin, external stringers, and internal
frames are utilized. The structure measures 83 in. high, by 168
in. at the base. The J-2 engine is attached at the small diameter
end of the thrust structure through the use of an A-356 aluminum
casting. The forward end of the thrust structure is attached to the
aft dome of the propellant tank assembly. Electro-mechanical and
mechanical systems equipment mounts on the inside and outside
surfaces of the thrust structure. Two doors provide access to thethrust structure interior.
EXTERNAL
T-EXTRUSION
ACCESS
ENGINE MOUNT
FITTING
(A-356 ALUMINUMMACHINED-CASTING)
;LE EXTRUSION
ATTACH RING
(THRUST STRUCTURE-
AFT DOME)
NTERNAL
Z-SECTION
FRAME
SEGMENT
C-H 13_-1
LONGITUDINAL TUNNELS.
The longitudinal tunnels (main tunnel and auxiliary tunnel) house
wiring, pressurization lines, and propellant dispersion system
shaped charges. The tunnel covers are made of 7075-T6 aluminum
alloy skin stiffened by internal ribs. These structures do not transmit
primary shell loads but act only as a fairing reacting to local
aerodynamic loading.
AERODYNAMIC FAIRING.
The aerodynamic fairing on the aft end of the interstage is a short
cylinder, 260 in. in diameter, made from eight 7075_T6 corrugated
aluminum alloy skin panels. This structure does not transmit pri-
mary shell loads but must maintain structural integrity when loaded
by local aerodynamic pressure.
PROPELLANT TANK CRYOGENIC INSULATION.
Another design requirement for the propellant tank and common
bulkhead is that they provide insulation for the LH 2 tank and
in the case of the common bulkhead that it prevent lox freezing
due to LH 2 temperatures during a ground hold. Figure 5-8 showsthe methods and materials used to provide the necessary insulation
for the LH z tank. Cross-sections show insulation details of theforward dome, aft dome, forward dome access door, aft dome and
common bulkhead area, and cylindrical section and aft dome area.
Generally, the insulating tile (a polyurethane foam and fiberglass
composite) bonds to the tank skin with an epoxy base adhesive
system. In certain areas shown, balsa pads are used in conjunction
with tile to line the tank walls with insulating materials. An insulat-
ing gap filler material consisting of a glass fiber and polyurethane
adhesive composite is used to fill gaps between tiles and around
fasteners. Liners consisting of woven glass impregnated with po-
lyurethane adhesive are bonded to the tile and balsa insulating
material. In critical areas around brackets and joints the liner is
covered by doublers of the same material.
ABLATIVE INSULATION.
Figure 5-9 shows the ablative coating material and the patterns
and depth to which it is applied to the external skin of the stage.
Figure 5-7
FORWARD DOME DEBRIS SHROUD.
A debris shroud surrounds the forward dome of the propellant
tank 40.125 in. above the equatorial plane of the dome. Eight nylon
cloth shroud segments 9 in. wide by 89 in. long are stitched together
with dacron threads and reinforced at their joint with eight nylon
cloth splice segments 4 in. by 9 in. The shroud segments are attached
to the dome with Velcro tape adhesively bonded to the dome.
RETROROCKET IMPINGEMENT CURTAIN.
A retrorocket impingement curtain is used to shield, at stage sepa-
ration, the area between the aft dome and aft skirt. The curtain
installation spans the area between the aft end of the aft skirt
and the engine thrust structurejunctionwith the aftdome. The
basic material used in the construction of the curtain is glass cloth.
Tape is used to seal openings in the curtain around boots, slots
in aluminum, and other openings.
PROPULSION.
Prime thrust for the S-IVB stage is provided by the main propulsion
system, which consists of a bipropellant J-2 rocket engine, fuel
system, oxidizer system, tank pressurization systems, engine chill-down systems, and a propellant utilization (PU) system. The J-2
engine, burning lox and LH 2, provides the thrust during the second
boost phase of flight to inject the S-IVB stage and payload into
orbit. In addition to supplying prime thrust, the J-2 engine provides
thrust-vector steering (pitch and yaw) for inflight course correction
during powered flight. Command signals from the IU guidance
and control system effect the flight steering by hydraulically gim-
balling the engine up to ±7 deg from the stage longitudinal axis.The PU system controls the mass of propellants loaded into the
stage by providing inputs to the propellant loading system.
The stage propellant tanks consist of a cylindrical section enclosed
by hemispherical bulkheads with an internal common bulkhead
dividing the structure into fuel and lox tanks. The tank capacity
is designed to satisfy an engine mixture ratio of 5 to 1 (oxidizer
5-8
Section V S-IVB Stage
S-IVB PROPELLANTTANKINSULATIONTANK r'-- JAMB BALSA-
ACCESS DOOR _ / INSULATING RING /ACCESS DOOR BALSA _ J t---JAMB SKIN /
INSULATING RING----_ _ J| INSULATING TILE /
WOVEN . L_I/ IGLASS /_ q"_-_l ._-
,_ INSULATING GAP FILLER_-,._
f¢ ....
WOVEN WOVEN GLASS DOUBLER _s I_--x,_GLASS
DOUBLE
ACCESS DOOR - JAMB JUNCTION
CRYOGENIC INSULATION DETAILS
AFT DOME - CYLINDRICALFORWARD DOME SECTION JUNCTION
SEGMENT SKIN_--_ INSULATING TILE _CRYOGENIC INSULATION DETAILS
wovEN _ /--S.,PLAPJO,NT / _;i_ \_
INSULATING TILE _ 1_INSULATING TILE I I AW
.Y.,CALT.LEA%,_.NER I_ _".ON0,NG0 COMMO "BULK HEAD i:;_ i::_:_"......
_Y E'T-HY LEN-E y/_/ /
TEREPHTHALATEAIFILM LINER (13) ; _._t_,_
(ALUMINIZED SIDE /FACING OUTWARD)
METALIZED MYLAR FILM /
EXTERNAL AFT DOME /INSULATION LINER j
(METALI '.ED SIDE /FACING OUTWARD)-_/
ENGINE,_E'
NG INSULATION
GLASS FIBERS
COMMON
NSTRUCTURAL
CRYOGENIC ADHESIVE
EXTERNAL INSULATION
BLANKET (13)
AFT ENVIRONMENTAL
CONTROL CHAMBER SEAL
AFT DOME EXTERNAL INSULATION AND
ENVIRONMENTAL CONTROL CHAMBER DETAILS
r--- INSULATING
TILE
f BALSA INSULATING
PAD
4.........-_CYLIN DRICALSECTION SKIN
--- AFT SKIRTATTACHING RING
......-----'--- CABLE SUPPORT
BRACKET (4)
DOME SKIN
_WOVEN
GLASS LINER
GLASS LINER
GLASS DOUBLER
_ALSA
INSULATING PAD
ANE
FOAM
AL WELD
(REF)
AFT DOME SKIN
COMMON BULKHEAD - AFT DOME JUNCTION
CRYOGENIC INSULATION DETAILS
Figure 5-8
C-H 14307
5-9
Section V S-IVB Stage
I
AUXILIARY
TUNNEL
-- LH 2 CHILLDOWNRETURN FAIRING
..............
::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::
NOTES:
LH 2 FILL & DRAIN
LINE FAIRING 7
f
ii i::iii i!
://:_::_}}i_i_::_i_::_i?iii;i::ii::::_?i::i::i::::.:i::::)_III
iiii!i!i!i_
ii_ii!_ii!iil
I. AREAS SHOWN BY THE FOLLOWING
SHADING HAVE ABLATIVE COATING
OF THE INDICATED THICKNESS.
!!_i!i!:::iii:_io.o12_o.oo_,-o.oo_1_N.0.020 ('_0.004, -0.005) IN.
0.030 + 0.005 IN.
0,035+0.005 IN.
ANTENNA (4)
TUNNEL
ULLAGE
ANTENNA (2)
LH 2 CHILLDOWNAPS MODULE (2) PUMP FAiRIN(
LOX VENT
IIII
2. AREAS SHOWN BY THE FOLLOWING
SHADING ARE SPECIALLY PREPARED
ENAMELED SURFACES WITH COATING
OF THE INDICATED THICKNESS.
_ 0.25 + 0.05 IN.
Figure 5-9
LH 2 NON-
PROPULSIVE
VENT (2)
_LH 2 FEED LiNEFAI RI NG
_i::i::i::i::iilJ_::!_..............,,;_¢+:e:.:_
iiii!ii::iiiiiii_iiii!iii!iii!iiiiiiii_iiiiiiii_i======================================================
-i::ii:: ii! iiiii:i
i!!ili!iiiiii_ii_iii!iiiiiiiiii_iiiiiiiiiiiiii_iii
i_i::iiiii_i::ili::ii::i::_i::i::i::i::!::i}ii:i:i:,i:,iiii:,iii:i%i_!:iiiii%iiiiiiil
CH-20131
to fuel by weight). The fill systems are sized to minimize countdown
time and also to be compatible with the other systems loading.
Lox design flow is 1000 gpm; LH z design flow is 3000 gpm. Initial
fill rates are slower to accomplish tank chilldown and to prevent
ullage pressure collapse. Final fill rates are also slower to providetank "topping" during the terminal count.
5-10
The propellant tank vent systems were designed to protect the tank
structure under all conditions of propellant tank loading, stage
power flight, and orbital venting. During loading, tank pressuresare maintained well below the normal tank prepressurization levels.
The negative pressure differential across the common bulkhead
between the fuel and oxidizer tanks is the limiting factor. This
35
34
36
33
\
I 2 3 4 5 6 7 8
22 LOX TURBINE BYPASS VALVE
23 MAIN FUEL VALVE
24 SEQUENCE VALVE (FUEL)
25 THRUST CHAMBER
26 LOX PUMP PRIMARY SEAL DRAIN
27 FUEL INLET MANIFOLD
28 START TANK RECHARGE LINE (BLOCKED)
29 EXHAUST MANIFOLD
30 HEAT EXCHANGER COLD HELIUM INLET
31 HEAT EXCHANGER
32 PRIMARY INSTRUMENTATION PACKAGE
33 LOX TURBOPUMP TURBINE
34 LOX TURBINE
35 ACCUMULATOR SUPPLY LINE
36 MIXTURE RATIO CONTROL VALVE
37 INTEGRAL START TANK
38 TANK SUPPORT AND FILL VALVE PACKAGE
39 FUEL PUMP DRAIN LINE
40 LOX PUMP DRAIN LINE
41 START TANK DISCHARGE HOSE
42 START TANK DISCHARGE VALVE
43 AUXILIARY INSTRUMENTATION PACKAGE
Figure 5-10
Section V S-IVB Stage
LOX INLET DUCT
THRUST CHAMBER INJECTOR
LOX TURBOPUMP DISCHARGE LINE
PURGE CONTROL VALVE
MAIN OXIDIZER VALVE
LOX BLEED VALVE
SEQUENCE VALVE (LOX)
LOX BOOTSTRAP LINE
FUEL INLET DUCT
ASI FUEL LINE
PNEUMATIC CONTROL PACKAGE
GAS GENERATOR CONTROL VALVE
FUEL BLEED VALVE
GAS GENERATOR COMBUSTOR
FUEL BOOTSTRAP LINE
FAST SHUTDOWN VALVE
ELECTRICAL CONTROL PACKAGE
FUEL TURBOPUMP
FUEL TURBOPUMP TURBINE
MAIN FUEL DUCT
LOX TURBINE BYPASS DUCT
37
36
CH.14395.|
5-1 1
Section V S-IVB Stage
ENGINE J-2033 PERFORMANCE TAG VALUES
PARAMETER
THRUST
SPECIFIC IMPULSE
CHAMBER PRESS
LOX FLOWRATE
FUEL FLOWRATE
MIXTURE RATIO
GAS GEN LOX FLOWRATE
GAS GEN FUEL FLOWRATE
LOX PUMP SPEED
FUEL PUMP SPEED
UNIT OF NOMINAL
MEAS TAG VALUES
LB 229,294
SEC 423.4
PSIA 770.6
LB/SEC 458.78
LB/SEC 82.78
O/F 5.542
LB/SEC 3.46
LB/SEC 3.55
RPM 8, 682
RPM 26,494
CH-14397-1
Figure 5-11
differential pressure is never allowed to exceed 3.0 psid for normalcontrolled operation.
All components in the propulsion system are designed to function
in an explosive atmosphere without providing a source of ignitionor electromagnetic interference. Potential leak sources have been
lowered and component weight reduced by the modular concept,
in which components such as pressure regulators, check valves and
solenoid valves are packaged in a single body, thus eliminating
many connections.
A temperature conditioning system is provided in both the LH 2
and lox feed systems to provide temperature stabilization at the
engine pump inlets to meet minimum net positive suction headrequirements.
VALUES
J-2 ENGINE SYSTEM.
The 225,000 lbf thrust, high performance lox and LH 2 J-2 engine(figure 5-10) powers the S-IVB stage/IU/payload during the second
boost phase of powered flight. The J-2 engine accelerates the
payload to orbital velocity during its 7 rain 30 sec (approximately)
burn. Lox, LH 2, and helium constitute the only fluids used on
the engine, because the extremely low operating temperature of
most of the engine components prevents the use of lubricants or
other fluids. The engine features a single tubular-wall bell-shaped
thrust chamber, and two independently driven turbopumps. Asingle gas generator, powered by lox and LH 2 bled off the main
LH 2 and lox turbopump discharge lines, drives both the lox andLH 2 turbopumps in series. A mixture ratio control value achieves
two mixture ratio levels by passing lox from the discharge side
of the turbopump to the inlet. A pneumatic control system provides
regulated helium for engine valve operations from a helium tank
mounted on the engine. An electrical control system, containing
solid-state logic elements, sequences the engine start and shutdown
operations. The S-IVB stage electrical power system supplies elec-
trical power for engine operation. An engine-mounted heat ex-
changer, located in the lox turbopump turbine exhaust duct, heats
helium from the S-IVB stage cold helium bottles for pressurizationof the lox tank. A bleed line from the thrust chamber fuel manifold
taps off GI-I 2 for LH 2 tank pressurization. The J-2 engine consists
of the following systems: propellant feed, pneumatic-electrical con-
trol, gas generator and exhaust, thrust chamber and gimbal, and
flight instrumentation. See figures 5-11 and 5-12 for J-2 engine
performance values.
5-12
Characteristics.
J-2 engine, serial number J2046, will fly on the SL-2 mission. The
engine measures 80175 in. in diameter and 133 in. in length. A
gimbal assembly and two hydraulic actuators attach the engine
to the S-IVB stage thrust structure. The gimbal assembly permits
the entire engine to gimbal for thrust vector control. Engine weight
with accessories is 3536 Ibm dry and 3697 Ibm wet. At burnout
the engine weighs3665 Ibm.
Acceleration and Velocity. Engine gimbal angular acceleration shall
not exceed 80 rev/sec2. During the boost phase of flight, before
engine ignition, the engine shall not be subjected to accelerations
greater than presented in figure 5-13.
Flight Load Limits. During the engine operation phase of flight,the following load conditions should not be exceeded.
a. The longitudinal and lateral inertia load limits as shown in figure5-14.
b. Aerodynamic moments during gimbal operation shall not exceed
the following and shall be such that when combined with all other
loads, other maximum limitation specifications in this manual shallnot be exceeded.
1. No. 1 or no. 2 gimbal axis (see figure 5-15); aerodynamic mo-ments of -----110,000 in.-lb.
2. X or Z engine axis (see figure 5-15); aerodynamic momentsof -+-210,000 in.-lb.
c. Maximum aerodynamic pressure on the thrust chamber shallnot exceed 500 lb/ft 2.
Gimballing and Vehicle Loads. The designed limit of forward accel-
eration with respect to engine gimbal angle is shown in figure 5-16.This limit occurs simultaneously with the lateral acceleration rela-
tionship to forward acceleration as shown in figure 5-14.
ENGINE UNITOPJ[_ MEAN _:> STDDEV(%IPARAMETER MEAS VALUE ENG-TO-EN(
THRUST
(I) ALTITUDE
(2) SEA LEVEL
SPECIFIC IMPULSE
(1) ALTITUDE
(2) SEA LEVEL
MIXTURE RATIO
RATED DURATION
LOX FLOWRATE
(PUMP INLET)
FUEL FLOWRATE
(PUMP INLET)
CHAMBER PRESS.
(INJECTOR END)
CHAMBER PRESS.
NOZZLE -
STAGNATION)
AREA EXPANSION
RATIO
Ib 225,000 [_
Ib 156,400
see 423.8
see 293.81
O,'f 5.50 _>
sec 500
Ib/sec 449.3
Ib/sec 81.68
psia 762.6
0 sla 702.2
I 27.12:1
i iI_ BASED ON ROCKETDYNE ACCEPTANCE TEST DATA
RATED CONDITIONS
I_ STD DEV (%)
RUN-TO-RUN
0.18 0.16
0.18 0.16
0.18 0.16
0.85 0.21
0.85 0.21
0.23 --
C-H 14396. I
Figure 5-12
GeneralPurgeRequlrements.The engine systems receive purges prior to loading propellants or
preconditioning the thrust chamber, to clear the systems of moisture
and/or gases, which would solidify or otherwise prove hazardous
when the engine hardware is chilled.
Turbopump and Gas Generator Purge Requirements. Fifteen minutes
prior to dropping propellants into the engine ducts the turbopump
and gas generator receive a 6-scfm (nominal flowrate) helium purgewith a temperature range of 50 to 200 ° F and an 82- to 125-psia
pressure range. Helium purges the fuel pump seal cavity, the fuel
turbopump turbine seal cavity, the gas generator fuel cavity, and
the lox turbopump turbine seal cavity. All four purge lines contain
check valves to prevent reverse flow of propellants into the purge
manifold. All except the fuel seal cavity purge line contain an orificedownstream of the check valves. The pneumatic control system
supplies the helium purge. See Pneumatic Control System foradditional information.
Oxidizer Dome, Gas Generator Oxidizer Injector, and Oxidizer In-
termediate Seal Purge. Approximately 15 min before dropping
propellants into the engine ducts, the oxidizer dome and GGoxidizer injector receive a 230-scfm (nominal flowrate) helium
purge. An electrical command from the component test heliumcontrol solenoid switch initiates the purge by opening the helium
control valve on the pneumatic control package. The control pack-
age regulates helium from the integral start tank to 400 ___25 psig.Helium enters the oxidizer dome through an orifice and a check
valve in the downstream side of the main lox valve housing and
escapes through the thrust chamber injector. The GG oxidizer
injector purge enters the gas generator through a check valve and
escapes through the turbine exhaust system. The oxidizer pump
intermediate seal purge is operative any time the pneumatic control
package is activated. The purge flowrate is 2600 to 2700 scimHelium purges the turbopump area between the pump housing
and the turbine inlet manifold. The three purges continue for 15
min and are terminated just before lox loading begins approxi-
mately 5 hr before liftoff. During flight, the J-2 engine start signalinitiates the three purges prior to lox entry into the lox dome and
gas generator. The oxidizer dome purge and GG oxidizer injector
purge continue for just over I sec when the mainstage control valve
J-2 ENGINE
COORDINATE AXIS DIAGRAM
1.0
0.9
0.8
0.7
o
0.6
0.5
_0.4
0.3
0.2
0.1
\\
\\
2 3 4 5 6 7 8
FORWARD LONGITUDINAL ACCELERATION (G)
Figure 5-13
10 11
C-H 14508-I
Section V S-IVB Stage
1.0
0.9
0.8
Z 0.7
O
0.6
_ 0.5
_ 0.4
0.3
\\
\
0.2
0.1
00 1 2 3 4 5 6 7 8 9 I0 11
FORWARD LONGITUDINAL ACCELERATION (G) C-H 14524-1
Figure 5-14
terminates the purges by closing the purge control valve. The
oxidizer intermediate seal purge continues throughout J-2 engine
operation. At J-2 engine cutoff, the mainstage control valve deener-gizes and permits the purge control valve to reinstate the oxidizer
dome and GG oxidizer injector purges, which terminate 1 sec after
the cutoff signal when the helium control valve closes and shuts
down the pneumatic control package operation. Oxidizer interme-diate seal purge ends at helium control valve closure.
Thrust Chamber Jacket Purge and Chilldown. The thrust chamber
jacket purge begins approximately 15 min before liftoff. Helium
at 50 psig and ambient temperature enters the engine fuel inlet
manifold and flows through the thrust chamber jacket cooling tubes
escaping through the thrust chamber injector. S-IVB pneumatic
console 433 supplies the purge through the aft umbilical panel
at 0.01 lbm/min. The thrust chamber purge valve switch opens
the supply valve. Open and closed indicators on the engine panel
monitor the supply valve position. This purge lasts for 5 min and
NO. 2
ACTUATOR (Y)
ATTACH
POINT ---_
POS IV \ +Z
NO. 2 \ / ,_..GIMRAL N/
POS III
cOXIDI ZER PUMP
_NO. 1
ACTUATOR (P)
ATTAC H/ t_ GIMBAL
POINT POS I / 'I CENTERLINE
+XG_i_B 1 FUEL PUMPL
OXIDIZER
PUMP POS II +y
Figure 5-15
C-H 14509-I
5-13
SectionV S-IVB Stage
then the thrust chamber chilldown operation begins. Cold helium
at approximately -320 ° F (min) and 1000 psig (max) enters the
S-IVB stage through the same umbilical connection as the jacket
purge and flows at 15 lbm/min into the thrust chamber jacket.
The S-IVB engine panel also controls this operation by the thrust
chamber chilldown switch, which opens the cold helium supply
valve. Position switches in the supply valve provide open or closed
signals to the thrust chamber chilldown indicator lights on theengine panel. Thrust chamber chilldown continues until the S-IB
ignition command at T-3 sec terminates the chilldown operation.
J-2 Engine Predicted Performance.
Predicted J-2 engine ignition will occur about 2 min 30 sec into
flight with a guidance cutoff occurring at about 10 min into flight.If however, the guidance cutoff does not occur at that time, the
LH 2 depletion cutoff sensors in the LH 2 tank are armed to preventpropellant starvation. Thrust decay causing TOPS to deactivate
will also initiate cutoff. See figure 5-17 for predicted J-2 engineperformance values.
\\
\\
\
J-2 Engine Operation.
Approximately 2 min 30 sec into flight, the S-IVB stage switch
selector issues the engine start signal to the electrical control pack-
age on the J-2 engine (figure 5-18). The electrical control package
performs all the sequencing functions for proper engine operations
and requires only dc power and start and cutoff commands for
operations. The engine start signal and an engine ready signal
initiate engine operation. An engine ready circuit monitors the
conditions and events that are necessary for engine start and issues
an engine ready signal when all pre-start conditions have been
met. Those prerequisites are: oxidizer turbine bypass valve open;connectors installed; helium control valve, ignition phase control
valve, start tank discharge valve control valve, ASI system, main-
stage control valve, and GG spark system all deenergized; and
absence of an engine running signal. The LVDC also issues an
engine ready bypass signal immediately prior to engine start. The
start command energizes spark exciters in the sequence controller
that provide electrical energy to spark plugs in the gas generator
and the augmented spark igniter system. A start tank dischargedelay timer (0.640 -- .030 sec) starts and the helium control valve
and ignition phase control valve energize simultaneously. 3000-psig
helium from the helium tank, in the integral start tank, flows
through the helium control valve into the pneumatic control pack-age. Roughing, primary, and control regulators reduce the pressure
to 400 -+- 25 psig for valve control. Helium from the primary
regulator outlet flows to continuously purge the lox turbine inter-
mediate seal during J-2 engine operations. Regulated helium from
the pneumatic control package flows into the jacket of the primary
flight instrumentation package, which serves as an accumulator
for the control pressure system. In the event of supply pressure
failure a check valve in the regulator outlet will prevent control
pressure loss and the accumulator would then supply a large enough
volume of helium to effect safe shutdown of the engine. The helium
also closes the fuel and lox bleed valves, and flows through the
purge control valve to purge the lox dome and GG oxidizer injector.
Helium flows through the normally open port of the mainstage
control valve to the main lox valve closing port, to the opening
ports of the purge control valve, and the lox turbine bypass valve.
Helium flows through the energized ignition phase control valve
(normally closed port) to open the main fuel valve and augmented
sparkigniter (ASI) oxidizer valve, and to the sequence valves on
the main fuel and lox valves. LH 2, under tank pressure, flows
through the turbopump and main fuel valve into the thrust chamber
fuel manifold. Bootstrap fuel tapped off the fuel manifold flows
to the ASI assembly. Lox, under tank pressure, flows through the
lox turbopump and ASI valve to the ASI assembly where the
0 1 2 3 4 5 6 7 8 9 10
FORWARD LONGITUDINAL ACCELERATION (G)C.H 14510-1
Figure 5-16
sparkplugs ignite the propellant. Fuel enters the fuel manifold and
flows through the down tubes to the return manifold and then
through the up tubes to the injector where it emerges in a gaseous
state. The fuel cools the thrust chamber as it flows through thejacket to the combustion chamber. Lox enters the oxidizer manifold
in the injector and flows through equally distributed nozzles where
it mixes with fuel and burns. The MFV sequence valve opens when
the MFV opens to 90 percent permitting helium flow to the closed
start tank discharge valve (STDV) control valve. The start tank
discharge timer expires after approximately 1.0 sec, and an ignitionphase timer (0.450 -+- 0.030 sec) starts. The STDV control valve
energizes, admitting helium to the STDV opening port. GH 2 from
the integral start tank flows through the series turbine drive systemaccelerating the fuel and oxidizer turbopumps to the required levels
to permit power buildup of the gas generator and to deliver propel-
lant for ASI ignition. The normally open lox turbine bypass valve
bypasses a percentage of the GH z from the oxidizer turbine during
start and later controls the relationship of fuel and oxidizer turbinespeed by an orifice in the valve gate.
Expiration of the ignition phase timer initiates: (1) command for
the sparks deenergized timer (3.30-+-0.20 see), (2) closing signal
ill 1_'I_1:1 I'J:i I_1 [l I I,l'J:i I
PARAMETER
AVERAGE THRUST (L8)
AVERAGE CHAMBER iNJECTOR PRESS. (psia)
AVERAGE EFFECT. THRUST COEF.
AVERAGE SPECIFIC iMPULSE (sec)
AVERAGE MIXTURE RATIO (O/F)
AVERAGE FUEL FLOWRATE (Ib/sec)
AVERAGE LOX FLOWRATE (Ib/sec)
MR SHIFT TIME
ENGINE CUTOFF
l*l:l;,l:l'lll
PRIOR TO
MR SHIFT
230,073
774.2
1.74
424.2
5.509
83.324
459.045
328. I
i'I:Ull Su
AFTER
MR SHI FT
193,655
659.2
1.72
428.2
4.786
78.162
374.068
448.1
TIME FROM S-IB OBCO (see)
Figure 5-17
CH.14314.
5-14
SectionVS-IVBStage
PREREQUISITES
OXIDIZER TURBINE BypASS VALVE OPEN
CONNECTORS INSTALLED _g_
HELIUM CONTROL VALVE DEENERGIZED
IGNitiON PHASE CONTROL VALVE DEENERGIZED
STAGE TANK DISCHARGE VALVE COhn ROL VALVE DEENERGIZED
MAINSTAGE CONTROL VALVE OEENERGIZED _.
AUGMENTED SPARK IGNITER SPARK SYSTEM DEENERGIZED,_ D
GAS GENERATOR SPARK SYSTEM DEENERGIZED =_
MAINSTAGE OK DEENERGIZED
IGNITION mE ENERGIZED'
CONTROL POWER ON
MAINSTAGE OK NO. I AND NO. 2 PRESSURE SWITCHES OPEN"_'=m-'_
ENGINE RUNNING C
LVDC ENGINE READY BYPASS
START COM.V.AND --_
ENGINE
START_ START TANK DISCHARGE .-I
ENGINE DELAY TIMER
(1.00 SEC)
ASI AND GAS GENERATOR J
SPARKS ON
VALVE CONTROL
EVENTS SEQUENCE
(_ HELIUM CONTROL VALVE OPENS. HELIUM FROM _NTEGRAL START TANK ENTERS PNEUMATIC CONTROLPACKAGE. ROUGHING, PRIMARY, AND CONTROL REGULATORS REDUCE 3000 FSIG INPUT TO 400 PSIG
" OUTPUT. LOX INTERMEDIATE SEAL PURGE BEGINS.
(_ FUEL AND LOX BLEED VALVES CLOSE (PROPELLANT RETURN FLOW TO TANKS STOPS). LOX fiLEED VALVEREMAINS OPEN TO LOX BOOTSTRAP LINE.
(_ PURGE CONTROL VALVE OPENS. LOX DOME AI 'D GAS GENERATOR LOX CHAMBER PURGE BEGtNS.
(_ IGNITION PHASE CONTROLVALVE ENERGIZES.
Q MAIN FUEL VALVE OPENS. AT 90% OPEN, MAtN FUEL VALVE MECHANICALLY OPENS THE SEQUENCE VALVE.CONTROL PRESSURE DIRECTED TO START TANK DISCHARGE VALVE CONTROL VALVE.
(_ ASI VALVE OPENS.
MAIN FUEL AND ASI PROPELLANT FLC_V EFFECTED BY LOX AND LH 2 TANK PRESSURES. FUEL FLOWS
THROUGH STATIC TURBOPUMP AND MAIN FUEL VALVE INTO THE FUEL MANIFOLD AND INTO THE AEI
ASSY. LOX FLOWS THROUGH STATIC TURBOPUMP AND ASI VALVE INTO THE ASI ASSY. TWO SPARK
PLUGS IGNITE THE PROPELLANT.
LL_ZX=_ _ IGN'TION
ASI SPAR K --="_DI_
START TANK DISCHARGE DELAY TIMER EXPIRES.
tGNtTION PHASE TIMER ENERGIZED (0.450+ 0.030 SEC).
(_ TANK DISCHARGE VALVE. FROM INTEGRAL START TANK STARTS FUELCONTROL VALVE OPENS START GH 2
TU_SOPUM?. GH 2 EXHAUSTS THROUGH LOX TURBINE BYPASS VALVE AND LOX TURBtNE tNTO EXHAUST
MANIFOLD.
IGNITION PHASE START
TIMER EXPIRESC(_MAND MAINSTAGE
T COMMAND
VALVE CONTROL
pOWER
(_ MAINSTAGE CONTROL VALVE OPENS.SPARKS DEENERGtZED TIMER ENERGIZED (3.30_+ 0.20 SEC)
PURGE CONTROL VALVE CLOSES. LOX DOME AND GAS GENERATOR PURGE STOPS.START TANK DISCHARGE VALVE CLOSES.
(_ MAIN LOX VALVE BEGtNS TO OPEN CONTROL PRESSURE AT FRE°STAGE OPENING PORT INITIATES VALVEI
OPENING. MAIN LOX FLOW TO THE LOX INJECTOR BEGINS. MAtN LOX VALVE SEQUENCE VALVE OPENS
AND DIRECTS CONTROL PRESSURE TO THE GAS GENERATOR CONTROL VALVE AND TO LOX TURBINE BYPASS VALVE •
LOX TURBINE BYPASS VALVE CLOSES PNEUMATICALLY.GAS GENERATOR CONTROL VALVE OPENS PNEUMATICALLY. FUEL AND LOX BOOTSTRAP PROPELLANTS
FLOW THROUGH CONTROL VALVE INTO THE COMBUSTOR (FUEL LEADS LOX). TWO SPARK PLUGS IGNITE
THE PROPELLANTS, WHICH EXHAUST THROUGH THE FUEL TURBINE TO THE LOX TURBINE.
_) MAIN LOX VALVE OPENS F_.LY. ORIFICED PRESSURE AT SECOND STAGE OPENING PORT PROVIDES BEOWOPENING OF LOX VALVE.
(_) GASGENERATOREXHAUSTFROMPUELTURBINEDUCTEDTHROUGHLOXTURB_NE. LOX AND FUEL TURBO-PUMPS ACCELERATE TO STEADY-STATE LEVEL.
IJilte TO ASSURE BALANCE IN FUEL AND LOX TUREOPUMP PERFORMANCE, A CALIBRATED NOZZLE
IS INSTALLED IN THE GATE OF THE LOX TURBINE BYPASS VALVE TO DUCT PART OF THE FUEL
TURBINE EXHAUST GASES TO THE EXHAUST MANtFOLD.
(_ OK PRESSURE SWITCHES ACTUATE WHEN LOX PRESSURE REACHES 500 __ 25 PSI • L/_I ENGINE NO. 1THRUST
LIGHT EXTINGUISHES.
NOte ENGINE ATTAINS 90_ THRUST APPROXIMATELY 3.6 SEC AFTER START COMMAND.
SPARKS DEENERGIZED TIMER EXPIRES. POWER REMOVED FROM ASt AND GAS GENERATOR SPARK PLUGS
COMBUSTION SUSTAINED BY CONSTANT FLOW OF PROPELLANTS INTO GAS GENERATOR AND ENGINE
COMBUSTION CHAMBER. r_
THRUST OK PRESSURE SW NO. I _'J-_ _ _ MAINSTAGE OK
THRUST OK PRESSURE SW NO. 2 _=_1_ _IA_ _ ENGINE
SPARKS DEENERGIZED TIMER EXPIRATION "--_1_ CUTOFF
ENGINE READY oL,'
HEAT EXCHANGER EXPANDS HELIUM FROM COLD HELIUM TANKS FOR INFLIGHT LOX TANK PRESSURIZATION. --GH2 fiLED FROM FUEL INJECTOR PROVIDES INFLIGHT PRESSURIZATION FOR LH 2 TANK.
(_ CONTROL VALVE BYPASSES LOX RACK TOTHE MIXTURE RATIO
THE LOX TURIK3PUMP INLET AND PROVIDES OPERATION AT5.5 OR 4.8 EMR LEVELS.
I
_llJliil lilJllli
SECONDS
IIIIJllll IIIIIII
PROPELLANT FLOW
[]
(:3
n
GH 2 FLC_
r-1
T
!
[] •
r_
E3r--
I
"1 ENABLED
L I
,,,,I,,, ,,,,l,,,,i,,,,I,,,, _J,l,,,_t0 I 2 3
_,,1_,,,'1
Ii
PUMP BUILDUP_.
MAIN LOX FLOW_
C,_S GENERATOR PROPELLANT FLOW _"
I
i ii J I itll
CH-14197-_
Figure 5-18 (Sheet 1 of 2) 5-15
Section V S-IVB Stage
J-2 ENGINE STARTI NTEGRAL HELIUM CONTROL
AND HYDROGEN_ @ .
._._ "'-'STARTTANK
VALVE (NC)
I_ FROM FUEL
_)iK FUEL TURBOPUMP-
[_:> PRE-STAGE
OPENING CONTROL
_::> SECOND STAGE
OPENING CONTROL
_ OUTER SHELL AND COVER SERVES
AS AN ACCUMULATOR FOR
PNEUMATIC CONTROL SYSTEM,
RETAINS ENOUGH VOLUME TO
EFFECT SAFE OPERATION AND
ENGINE SHUTDOWN IF LOSS OF
HIGH P_ESSURE SUPPLY OCCURS.
OPENING PRESS.
CLOSING PRESS.
SUPPLY HELIUM
SIGNAL
• PROBE
ASSY
FUEL EXHAUST
MANIFOLD-_ MANIFOLD
THRUST
5-16 Figure 5-18 (Sheet 2 of 2)
for the start tank discharge valve, and (3) mainstage command.
Energizing the mainstage control valve vents pressure from themain oxidizer valve (MOV) closing port, closes the purge control
valve terminating the lox dome and GG oxidizer injector purges,vents the lox turbine bypass valve opening pressure, and applies
opening pressure to the first- and second-stage opening port ofthe MOV. (An orifice in the second-stage opening line, series orifices
in the actuator closing line, an orifice in the actuator, and an orificed
check valve provide a controlled ramp opening of the MOV.) The
MOV sequence valve opens with the MOV and supplies helium
pressure through orifices to close the lox turbine bypass valve and
to open the gas generator control valve. Two spark plugs in the
gas generator ignite the lox and LH 2 admitted by the GG control
valve. Hot-gas products of combustion pass through the fuel turbine
and through the exhaust duct to the lox turbine accelerating the
turbopump causing increased propellant flow. The turbine exhaust
gases exit through a heat exchanger and into an engine exhaust
manifold. The propellant enters the thrust chamber where main
propellant ignition occurs. As turbopumps accelerate to operational
speeds, oxidizer injection pressure increases actuating two thrust
OK pressure switches. Either of the two TOPS, which actuate at
500_30 psi, will issue a mainstage OK signal that must be present
before the sparks deenergized timer expires or the sequencer will
automatically issue a cutoff command. Expiration of the sparks
deenergized timer (3.30___0.20 sec) also turns off the ASI and GG
spark exciters. The mainstage OK signal extinguishes the L/V
ENGINE 1 light on the MDC, indicating that the engine has
attained 90-percent thrust. During engine operation, GH 2 tapped
off the fuel injection manifold maintains LH z tank pressure. Tur-
Section IV S-IB Stage
bine exhaust gases flowing through the heat exchanger, heats he-lium from the cold helium bottles for lox tank inflight pressuriza-
tion. A two-position mixture ratio control valve (MRCV) bypasses
lox from the oxidizer pump outlet to the pump inlet and allows
the engine to operate at either of two fixed mixture ratios. TheMRCV is commanded to the low EMR position prior to engine
start and to the high position after 90% thrust is achieved.
J-2 Engine Cutoff.Guidance cutoff of the J-2 engine occurs at about I0 rain into
flight. This command will be issued through the switch selector
to the sequence controller in the J-2 engine electrical control pack-
age. The sequence controller simultaneously deenergizes the main-
stage control and ignition phase control valves and energizes thehelium control deenergize timer (figure 5-19). Opening control
pressure vents from both the first- and second-stage main oxidizervalve opening actuators through the normally closed port of the
mainstage control valve while opening control pressure from the
augmented spark igniter oxidizer valve and the main fuel valvevents through the normally closed port of the ignition phase control
valve. Helium control system pressure is routed from the normally
open port of the mainstage control valve to the closing actuatorof the main oxidizer valve and to the opening control port of the
purge control valve. The purge control valve opens allowing helium
control pressure to flow to a check valve in the oxidizer dome
purge line and to a check valve in the gas generator oxidizer purgeline. Both of these purges will flow when thrust chamber and gas
generator chamber pressures decay below the level of control system
J-2 ENGINECUTOFF
O
(_ CUTOFF COMMAND DEENERGIZES MAINSTAGE AND IGNITION PHASE CONTROL VALVES,AND ENERGIZES HELIUM CONTROL DEENERGIZE TIMER
(_ MAIN OXIDIZER VALVE CLOSES PNEUMATICALLY AND TERMINATES LOX FLOW TO ENGINE.--1ST AND 2ND STAGE OPENING PRESS. VENTS THROUGH N.C. SIDE OF MAINSTAGE
CONTROL VALVE. LOX SEQUENCE VALVE MECHANICALLY CLOSES.
(_) ASI VALVE CLOSES PNEUMATICALLY TERMINATING LOX FLOW TO ASI ASSY. OPENING --PRESS. VENTS THROUGH N.C. SIDE OF IGNITION PHASE VALVE.
(_) VALVE CLOSES PNEUMATICALLY TERMINATING FUEL FLOW TO THE THRUSTMAIN FUEL
CHAMBER AND ASI ASSY. OPENING PRESS. VENTS THROUGH N.C. SIDE OF IGNITION
PHASE VALVE. FUEL SEQUENCE VALVE CLOSES MECHANICALLY.
(_ FAST SHUTDOWN VALVE OPENS PNEUMATICALLY AND RAPIDLY VENTS GG CONTROLVALVE ACTUATOR, LOX TURBINE BYPASS VALVE ACTUATOR, AND LOX SEQUENCE VALVE.
Q CONTROL VALVE TERMINATING LOX FLOW.GG LOX CLOSES,
GG FUEL CONTROL.VALVE CLOSES TERMINATING FUEL FLOW.
(_ AND LOX TURBOPUMPS SPEED DECAYS.FUEL
(_) CONTROL VALVE OPENS PNEUMATICALLY. OXIDIZER DOME AND GG OXIDIZER --PURGE
PURGES BEGIN WHEN ENGINE OPERATING PRESSURES DECAY BELOW PURGE PRESSURE.
(_) DECAYING LOX INJECTOR PRESS. CAUSES THRUST OK PRESSURE SWITCHES TO DEENERGIZE.
(_ 1 LIGHT ILLUMINATES ON THE CM MDC INDICATING J-2 ENGINE CUTOFF.L_/ ENGINE
(_) TURBINE BYPASS VALVE OPENS PNEUMATICALLY. CLOSING PRESS. VENTS THROUGH --LOX
FAST SHUTDOWN VALVE.
(_ CONTROL DEENERGIZE TIMER EXPIRES AND CLOSES HELIUM CONTROL VALVE TOHELIUM
SHUT OFF CONTROL PRESS.
(_ CLOSING PRESS. TO LOX AND FUEL BLEED VALVES VENTS BACK INTO PNEUMATIC CON-TROL PACKAGE, THROUGH IGNITION PHASE CONTROL VALVE N.C. PORT ( c ), INTOTHE MAIN FUEL VALVE CLOSING PRESS. LINE, AND THROUGH BLEED ORIFICE AND LOW-
PRESS. RELIEF VALVE. BLEED VALVES OPEN AND VENT PROPELLANT FEED LINES BACK
INTO THEIR RESPECTIVE TANKS.
(_ INTERMEDIATE SEAL PURGE STOPS.LOX
0.2
I
I--1
-------t
SECONDS
0.4 0.6
I I
i
I
-- (OPENS WITHIN TEN SECONDS)
Figure 5-19 (Sheet 1 of 2)
0.8 1.0
CH-14303-1
5-17
SectionVS-IVB Stage
HELIUM .."_
LOX IImmmmmmmmi INLET""
GH 2
GAS GENERATOR
EXHAUST
CONTROL PRESS.
OPENING PRESS.
CLOSING PRESS.
SUPPLY HELIUM
SIGNAL
GMENT SPARK
IIIIIIIIInlIIIIUlItlUlIlUllUlUlII I IGNITER VALVE
I ASSY (NC)I
l THRUST OK
PRESSURE
SWITCH
i LH TANKPRTSSURIZATION
FUEL
MANIF
THRUST
CHAMBER_
HELIUM
OUTLET
PROBE
ASSY
CH.14304.1
5-18 Figure 5-19 (Sheet 2 of 2)
pressure. The normally open port of the ignition phase control
valve routes helium to the closing actuator of the augmented spark
igniter oxidizer valve and the main fuel valve and to the opening
control port of the pressure-actuated fast shutdown valve which
actuates, allowing the gas generator control valve opening control
pressure to vent rapidly. All valves, except the augmented spark
igniter oxidizer valve and the oxidizer turbine bypass valve, are
spring-loaded to the closed position and start to close as soon as
opening pressure is vented. Combustion pressure in the gas genera-
tor assists the spring in closing of the gas generator" control valve.
The oxidizer turbine bypass valve closing control pressure also vents
through an orifice and the pressure-actuated fast shutdown valve.
The oxidizer turbine bypass valve is spring-loaded to the normally
open position and starts to open as closing pressure is vented. The
normally open port of the mainstage control valve also supplies
opening control pressure to the oxidizer turbine bypass valve to
assist the spring in opening the valve.
Decaying lox injector pressure causes the thrust OK pressure
switches to deactuate. L/V ENGINE 1 light on the main displayconsole in the command module illuminates at TOPS deactuation
providing a visual indication of J-2 engine thrust decay. The light
remains on until CSM/launch vehicle separation. Expiration of
the helium control deenergize timer causes the helium control valve
to deenergize, closing the valve and venting control system pressure
through the oxidizer dome and gas generator oxidizer purges. As
the control system pressure is vented, the normally closed purgecontrol valve actuates closed and the purges stop. Pressure is now
locked up in the system between the check valve in the pneumatic
control package, through the normally open port of the mainstage
and ignition phase control valves, the pneumatic accumulator, and
to the propellant bleed valve control ports holding the bleed valves
closed. The pressure in this system is bled off through an accumula-tor bleed orifice located in the line between the closing actuator
of the main fuel valve and the normally open port of ignition
phase control valve. As this pressure decays, the propellant bleed
valves open by spring pressure and the cutoff sequence is complete.
During the transient period, from the cutoff signal to zero thrust,
the engine will consume approximately seven gal of lox and 47gal of LH 2. Thrust will decay to 5% within 0.4 sec after receipt
of the cutoff signal. LH z and lox turbopump speeds will decay
to zero rpm in 10 and 4 sec, respectively.
Thrust Chamber and Gimbal System.
The thrust chamber and gimbal system consists of a thrust chamber,
where propellants burn to create thrust; the augmented sparkigniter, which ignites the propellants; and the gimbal assembly,
which allows the thrust chamber to correct pitch and yaw attitude
errors.
Thrust Chamber. The thrust chamber, consisting of a body and
an injector, receives liquid propellants under turbopump pressure,
converts them to a gaseous state, mixes them, and burns them
imparting a high velocity to the expelled combustion gases produc-
ing thrust for vehicle propulsion.
Thrust Chamber Injector. The concentric-orificed, porous-faced
thrust chamber injector (figure 5-20) atomizes and mixes propellants
to produce the most efficient combustion. Oxidizer ports are elec-
trically discharge machined to form part of the injector. Threaded
fuel nozzles install over the oxidizer ports to form the concentric
orifices. The injector face, formed from a sintered metallic material,
is welded at its outside and inside edges to the injector body. Each
fuel nozzle is swaged to the injector face. An oxidizer inlet elbow,
integral with the dome and injector assembly, admits lox from
the turbopump and injects it through the oxidizer ports into thethrust chamber combustion area. Fuel enters the injector from the
upper fuel manifold and flows through orifices concentric with
Section V S-IVB Stage
oxidizer orifices. Approximately 3 to 4 percent of the fuel flows
through the sintered injector face to cool the injector. Combustion
zone pressure acts upon the injector face area producing the thrustforce, which is transmitted through the gimbal to the vehicle struc-
ture.
Thrust Chamber Body. The tubular-walled, bell-shaped thrust
chamber, consists of a cylindrical section where combustion occurs,
a narrowing throat section, and an expansion section. The thrust
chamber body is constructed of longitudinal stainless steel tubes
brazed together with bands around the tubes for external stiffening.
A fuel inlet manifold on the engine bell admits fuel to 180 down
tubes that carry the fuel to the return manifold at the base ofthe thrust chamber. From the return manifold, 360 tubes carry
fuel to the thrust chamber injector. Fuel flowing through the thrust
chamber jacket cools the thrust chamber and at the same time
absorbs heat converting from LH 2 to GH 2 for injection into the
combustion chamber. Propellants burn in the combustion chamber
creating large volumes of gases that are forced to exit the combus-
tion chamber through the narrow throat area and out the drainage
nozzle producing thrust to propel the vehicle.
Gimbal Assembly. The gimbal is a universal joint consisting of a
spherical, socket-type bearing with a teflon-fiberglass composition
coating to provide a dry low-friction bearing surface. The gimbal
assembly, installed on the thrust chamber dome, attaches the engineto the thrust structure and transmits engine thrust to the vehicle.
A boot with a bellows configuration made of silicone-impregnated
fiberglass material protects the gimbal assembly from dust, water,
and other foreign matter without interferring with gimballing. Two
hydraulic actuators, attached to the engine and thrust structure,
provide the force to gimbal the engine +---7 deg for thrust vector
control. See Flight Control for information on the hydraulic system.
Augmented Spark Igniter Assembly. The augmented spark igniter
assembly is chamber mounted in the thrust chamber injector. Itconsists of a fuel and lox manifold, an injector assembly, and two
tHRUSTCHAMBERINJECTOR
', POROUS
_E, PACE_A _ 0 XIDI ZER
YI _ W POST
_- IZER\
POSTS " \
POROUS _ AUGMENTED SPARK
FACE IGNITER CHAMBERC-H 14322
Figure 5-20
5-19
SectionV S-IVB Stage
BEARING
NO.
LOCATION VIEW
LABYRINTH
SEAL
PISTON
BALANCE
RING
DIFFUSER
VANES
BEARING
NO. 2
LOW PRESSURE
SEAL
PRIMARY
SEAL
AVERAGE ENGINE PERFORMANCE
PARAMETER
PUMPENG INLET PRESS.
DISCHARGE PRESS.
DEVELOPED HEAD
VOLUMETRIC FLOWRATE
WT. FLOWRATE
HORSEPOWE R
SPEED
TURBINE
INLET PRESS. (TOTAL)
OUTLET PRESS.
INLET TEMP
OUTLET TEMP
MEAN
30.00 PSIA
1,224 PSIA
37,517.2 FT
8,414 GPM
82.486 LB/SEC
7,739.13 BHP
26,702.0 RPM
633.61PSIA
87.08 PSIA
1200.0 ° F
769,9 ° F
STD DEVIATION (%)
ENGINE RUN
TO TO
ENGINE RUN
0.;2 o._30.71 0.43
0.18 0.16
0.18 0.16
2.04 0.58
1.0 0.17
1.52 1.51
1.49 0.36
1.36 1.73
OMNI
SEAL
INLET
f%
_m_
f"
L
FIRST STAGE ,.j
TURBINE WHEEL /
SECOND STAGE /
TURBINE WHEEL
INDUCER NUT
_ _-" INDUCER
BEARING
SUPPORT
ASSY
7-STAGE
ROTOR ASSY
STATORASSY
COVER
SEAL
SECONDARY
SEAL
OUTLET
POLYURETHANE
FOAM
TURBINE
MANIFOLD
C-H 14224-1
5-20 Figure 5-21
spark igniters and cable assemblies. The ASI assembly receives
the initial flow of lox and LH 2 and ignites the propellants by
discharging electrical energy through the two spark plugs. At engine
start, spark exciters in the electrical control package transform 28
Vdc into 27,000-V (±3, 000) pulses that discharge across the spark
plug gap at 40 sparks per sec (min). Hermetically sealed transmis-
sion cabling and connections are pressurized with GN 2 to ensure
operation at high altitudes.
Propellant Feed System.
The propellant feed system consists of a fuel turbopump, oxidizer
turbopump, main lox valve, main fuel valve, augmented spark
igniter oxidizer valve, mixture ratio control valve, propellant bleed
valves, and propellant feed ducts. This system transfers and controls
propellant flow from the stage tanks to the thrust chamber and
the gas generator.
Fuel Turbopump. The fuel turbopump, a turbine driven, axial-flow
pumping unit consisting of an inducer, a seven-stage rotor, and
a stator assembly, increases the pressure and flowrate of LH 2
entering the thrust chamber. See figure 5-21. The pump is self-
lubricated and self-balanced with LH 2. The high speed, two-stage
turbine, driven by hot gas from the gas generator, drives the inducer
and the one-piece seven-stage rotor assembly. Gas enters the turbine
inlet manifold and passes through nozzles where it is expanded
and directed at high velocity through the first-stage turbine wheel
and then through stator blades that redirect the gases through the
second-stage turbine wheel. Exhaust ducts direct the gases to the
oxidizer turbopump turbine and also to the oxidizer turbine bypass
valve. Three dynamic seals in series prevent LH 2 and turbine gases
from mixing. The inducer increases the LH 2 inlet pressure at the
pump. Each stage of the seven-stage rotor contributes to the buildupof pressure which forces the fuel through diffuser vanes and outlet
volute into the high pressure duct. A self-compensating balance
piston absorbs axial thrust loads developed by the rotor. LH 2
lubricates and cools the two ball bearings on the rotor shaft. A
magnetic pickup senses pump speed as 12 equally spaced slots
in the rotor assembly interrupt the magnetic field. Temperaturemeasurements of the fuel turbine inlet and the fuel pump discharge,
a fuel pump discharge pressure measurement, and the pump speed
measurements are telemetered to ground receiving stations for
recording and postflight evaluation. See J-2 Engine Measuring.
Oxidizer Turbopump. The single-stage, direct-turbine-drive centrif-
ugal pump delivers lox to the thrust chamber at increased pressure
and flowrates necessary for satisfactory J-2 engine operations. See
figure 5-22. The pump is self-lubricated and self-cooled with lox.
A high speed, two-stage turbine, driven by exhaust gases from the
fuel turbine, provides power for lox turbopump operation. One
static and two dynamic seals in series prevent lox and turbine gases
from mixing. Turbine exhaust gases enter the oxidizer turbopump
turbine manifold and flow through nozzles and the first-stage
turbine wheel. Stator blades redirect the gases, which pass through
the second-stage turbine wheel and then exhaust through turbine
exhaust ducting, a heat exchanger, and through the thrust chamber.
A pump shaft transmits the turbine power to the inducer and
impeller. Lox enters the turbopump through the inducer, which
increases pump inlet pressure, and flows through the impeller into
the outlet volute. Passages from the outlet volute permit lox flow
to the two pump shaft ball bearings for lubrication. A screen filters
lox entering the bearing area. Seven measurements of turbopump
operation are taken during flight. The turbine inlet and outlet
temperatures and pressures, the pump discharge temperature and
pressures, and the pump speed are telemetered to ground receiving
stations and recorded for postflight evaluation. See J-2 Engine
Measuring. A magnetic pickup located behind the turbine manifold
Section V S-IVB Stage
senses pump speed as 12 equally spaced slots in the rotor assembly
rotate through the magnetic field. An accessory drive adapter on
the oxidizer turbopump drives the hydraulic pump for the two
engine actuators. The drive was designed for 30 hp extraction at
mainstage speeds. Engine performance balance is based on extrac-
tion of 15 hp. The 15 hp value was selected based on nominal
power requirements to gimbal the engine at model specificationlimits.
Main Oxidizer Value. The main oxidizer value (MOV) is a butter-
fly-type valve, spring-loaded to the closed position, pneumatically
operated to the open position, and pneumatically assisted to the
closed position. The MOV, installed between the high pressure
duct and the thrust chamber injector oxidizer inlet, controls the
lox flow into the thrust chamber. Pneumatic pressure from the
normally open port of the mainstage control valve plus spring
pressure maintains the valve closed. During the engine start se-
quence, the mainstage control valve opens and applies pressureto the first and second stage opening control ports on the MOV
actuator. Pressure acting against a small surface in the first stage
actuator plus helium flowing through an orifice and acting on a
large surface piston, which actuates the valve gate, provides a ramp
opening of the valve. Pressure exhausting from the closing actuator
exits through an orifice in the actuator housing and through an
orifice-check valve to help accomplish the ramp opening. The valve
opens in two stages. First stage actuator pressure will start opening
the valve 50 "+ 20 msec after the mainstage command to the
mainstage control valve. Within 50 -- 25 msec the valve will open
approximately 14 ± 2 deg. After 610 -+- 70 msec the second stage
pressure begins to open the MOV, and after 1825 -+- 75 msec the
valve will be fully open. A sequence valve, installed on the MOV
and actuated when the MOV opens, permits helium flow to the
gas generator control valve and to the closing port of the lox turbine
bypass valve. A spring closes the sequence valve when the opening
pressure is vented. A position switch assembly operated by the
gate shaft provides signals to the ESE concerning the gate positions.
The switch contains a potentiometer that senses valve position and
provides a corresponding input to the telemetry system. Open and
closed indicator switches operate the MAIN OXIDIZER VALVEOPEN and MAIN OXIDIZER VALVE CLOSED indicators on
the S-IVB engine panel during checkout and prelaunch operations.
During flight, signals from the MOV open position indicator switch
will be telemetered to ground stations and monitored on the engine
panel. The closed position indicator and valve position potentio-
meter signals will be telemetered to a ground receiving station and
recorded for postflight evaluation. See J-2 Engine Measuring for
additional information. A 5-Vdc reference voltage is supplied to
the position potentiometer and 28 Vdc is supplied to the indicatorswitches.
Main Fuel Valve. The main fuel valve is a butterfly-type valve,
spring-loaded to the closed position, pneumatically operated to the
Open position, and pneumatically assisted to the closed position.The MFV, installed between the fuel turbopump high pressure
discharge duct and the fuel inlet manifold, controls LH 2 flow tothe thrust chamber and to the ASI assembly. The ignition phase
control valve on the pneumatic control package opens the MFV
during engine start. When pneumatic pressure has opened the valve
to 90 percent, the actuator mechanically opens a sequence controlvalve mounted on the MFV acutator. The sequence valve permits
helium flow also from the ignition phase valve to the start tank
discharge valve control valve. The MFV and sequence valve ar-
rangement assures that the MFV will be open before the STDV
opens to deliver GH 2 to start the fuel turbopump. A position switch
assembly operated by the gate shaft provides analog position signals
of the gate and discrete signals corresponding to the open or closed
position of the gate. MAIN FUEL VALVE OPENED and MAIN
FUEL VALVE CLOSED indicators on the S-IVB engine panel
5-21
Section V S-IVB Stage
J-2 ENGNE LOX TURBOPUMP
TURBINEINLET
IMPELLER
ANT_-VORTEX RING*'=_
NOZZLE
BLADES
TURBINEWHEEL
IST STA(_E
TURBINE
WHEEL_
PUMPINLET
MfXTURE RATIOCONTROLVALVE RETURN
IMPELLERINLETSEAL
PUMPDISCHARGE
PARAMETER
OXIDIZER PUMP
ENG. INLET PRESS.
PUMP DISCHARGE PRESS.
DEVELOPED HEAD
_ VOLUMETRIC FLOWRATEHORSEPOWER
J_ SPEED
WT. PLOWRATE
OXIDIZER TURBINE
INLET PRESS. (TOTAL)
OUTLET PRESS. (STATIC)
INLET TEMP
OUTLET TEMP
AUX. PWR. AVAILABLE
(30 BHP MAX.)
MEAN
39.00 PSIA
1,080 PSIA
2,116.9 FT
2,907.4 GPM
2,201.91 EHP
8, _/1.9
485.55 LB/SEC
85.9 PSIA
32.5 PSIA769.86 ° F
612.15° F
15.0 BHP
STANDARD DEVIATIONS %
ENG-TO-ENG RUN-TO-RUN
0.63 0.33
0.58 0.41
0.17 0.16
0.60 0.52
0.54 0.14
0.17 0.16
1.72 0.37
1.56 0.43
1.36 1.73
1.41 1.04
EXCLUDES NOMINAL MIXTURE RATIO CONTROL VALVE BYPASS FLOW
INCLUDES NOMINAL MIXTURE RATIO CQNTROL VALVE BYPASS FLOW
5-22Figure 5-22
MIXTURE RATIO
CONTROLVALVE SUPPLY
Note
TURBOPUMP ROTATED120° CCW FROM PLIGHT
ATTITUDE
CH - 14201-2
monitor the valve positions through hardwire connections during
checkout and prelaunch operations. During flight, signals from the
MFV opened switch are telemetered to the ground, and, in addition
to being recorded, they are monitored by the MAIN FUEL VALVE
OPENED indicator on the S-IVB engine panel. The MFV closed
signal and the continuous position signals are telemetered to the
ground and recorded for postfiight evaluation. A 5-Vdc reference
voltage input to the position potentiometer and 28 Vdc to the
indicator switches is stage supplied. Based on Rocketdyne test
results, the MFV will begin opening 60±30 msec after the control
signal to the ignition phase valve. The valve requires an additional
110_50 msec to fully open. At engine cutoff, the MFV will beginclosing 90"+-25 msec after the cutoff signal issuance, and will be
completely closed 225-*-25 msec later.
Augmented Spark Igniter Oxidizer Valve. The normally closed,
pneumatically operated, poppet-type augmented spark igniter oxi-
dizer valve controls lox flow to the ASI assembly during J-2 engine
start sequence. The ignition phase control valve pneumatically
opens the AS! valve, and lox under tank pressure flows throughthe valve and into the ASI assembly where it mixes with LH 2.
The ASI valve mounts in the MLV gate housing and receives lox
through a port just upstream of the MLV gate. A position indicatorprovides an ASI LOX VALVE OPENED indication to the S-IVB
engine panel through hardwire connections during checkout. Dur-
ing flight the indication is telemetered back to ground stations andrecorded.
Misture Ratio Control Valve. The mixture ratio control valve allows
the engine to operate at either one of two fixed mixture ratios
to achieve maximum vehicle performance. The valve changes mix-
ture ratio by routing a portion of the oxidizer flow from the oxidizer
turbopump outlet back to the pump impeller inlet. The valve has
an actuator assembly and a gate assembly. The actuator is two-posi-
tion, electro-pneumatic and is spring-loaded to keep it in the highengine mixture ratio position (valve closed). Pneumatic pressure
is directed to the actuator piston by a three-way pneumatic controlvalve that is energized by a stage signal. The gate assembly consists
of a rotating sleeve within a stationary outer sleeve. Each sleeve
has three elongated holes; by rotating the inner sleeve (valve gate)the holes are alined or misalined, to control the amount of oxidizer
flow through the valve. The valve position indicator is mounted
on the valve shaft and consists of a rotary-motion, variable resistor
and open and close position switches.
The mixture ratio control valve has two distinct stops, to allow
engine operation at engine mixture ratios of either 5.5:1 or 4.8:1
(lox to LH 2 by weight). Pneumatic pressure is supplied to the valve
from the engine pneumatic system when the engine helium control
valve is energized. At a preselected time during engine operation,
a control signal, supplied by the stage, energizes the solenoid control
valve. Energizing the solenoid control valve allows pneumatic pres-
sure to enter the valve and apply force to the actuator piston,
to overcome the spring tension and move the piston in the direction
to rotate the gate to the low engine mixture ratio position (valve
open). Opening the valve results in a reduced oxidizer flow to the
thrust chamber. If either the pneumatic pressure or the electrical
command is lost, the valve will move to the high engine mixture
ratio position (valve dosed). The position indicator arm rotates
with the gate shaft, to remotely indicate valve position.
Propellant Bleed Valves. A propellant bleed valve in the lox systemand one in the LH 2 system bleeds trapped gases in the systems
back to their respective tanks. During chilldown operations, the
chilldown pumps circulate fuel and lox through the respective
systems. Propellants flow through the inlet ducts and turbopumps
Section V S-IVB Stage
and return to the stage tanks through the bleed valves. The valves
are poppet-type, spring-loaded to the open position and pressure
actuated to the closed position. The oxizider bleed valve is mounted
on the lox bootstrap line located on the oxidizer high pressure
duct just upstream from the MOV. The lox bleed valve has an
inlet port, two outlet ports, and an actuation port. Lox flows through
the valve to a return line to the lox tank and through the lox
bootstrap line to the gas generator. When actuated, the lox tank
return line port closes and the bootstrap port remains open. The
fuel bleed valve is mounted on the LH 2 bootstrap line at the fuel
turbopump outlet. This valve has an actuator port, an LH 2 inlet
port, and one LH 2 outlet port for fuel return to the LH 2 tank.At engine start, pressure from the pneumatic control package actu-
ates both bleed valves terminating propellant flow back to the tanks.
A position indicator in each valve feeds a 28-Vdc BLEED VALVE
CLOSED signal to the respective indicators on the S-IVB engine
panel during checkout operation. In flight, these signals are teleme-
tered to ground stations as events of valve actuation and are rercorded for evaluation.
Propellant lnlet Ducts. The fuel and lox inlet ducts convey the
propellants from the stage tanks to the fuel and lox turbopumps.Fhe ducts employ flexible bellows sections to permit freedom of
movement during engine gimballing. Bipod clevis assemblies stabi-
lize the bellows convolutions allowing maximum engine gimballing
without collapsing the bellows. A vacuum jacket insulates the fuelinlet duct to reduce boiloff of LH 2.
Gas Generator and Exhaust System.
The gas generator and exhaust system consists of a gas generator,
which supplies the hot gases to drive the turbopumps; turbine
exhaust ducts, which transfer the exhaust gases from the fuel turbineto oxidizer turbine and to the thrust chamber exhaust manifold;
the heat exchanger, which expands cold helium for lox tank inflight
pressurization; and the turbine bypass valve, which allows a portion
of the fuel turbine exhaust gases to bypass the oxidizer turbopumpturbine.
Gas Generator Assembly. The gas generator, which produces the
hot gases to drive the oxidizer and fuel turbines, consists of a
combustor containing two spark plugs, a control valve containing
oxidizer and fuel poppets, and an injector assembly. When engine
start is initiated, spark exciters in the electrical control package
are energized providing energy to the spark plugs in the gas genera-
tor combustor. Propellants flow through the open poppets of the
control valve to the injector assembly and into the combustor where
they are mixed and burned, resulting in hot gases that pass throughthe combustor outlet and are directed to the fuel turbine and then
to the oxidizer turbine. See figure 5-23 for gas generator charac-teristics.
Gas Generator Control Valve. The gas generator control valve is
a pneumatically operated, spring-loaded to the closed position,
poppet valve. The oxidizer and fuel poppets are mechanically linked
by an actuator. The purpose of the gas generator control valve
is to control the flow of propellants through the gas generator
injector. When the mainstage signal is received, pneumatic pressure
is applied against the gas generator control valve actuator assembly
which moves the piston and opens the fuel poppet. During the
fuel poppet opening, an actuator contacts the piston that opens
the oxidizer poppet. LH 2 and lox from the bootstrap lines flow
through the control valve into the combustion chamber. Orifices
in the bootstrap lines control the propellant flowrate to the gas
5-23
Section V S-IVB Sta
GAS GENERATOR CHARACTERISTICS
PARAMETER
CHAMBER PRESS.
(INJECTOR END)
OXIDIZER FLOWRATE
FUEL FLOWRATE
OUTLET TEMP
MEAN
654.7 psla
3.4 Ib/se¢
3.62 I b/see
1,200°F [_
STD DEV (%)
ENG-TO-ENG RUN-TO-RUN
1.46 0.51
1.08 0.51
1.08 0.51
[_:;> RATED CONDITION
Figure 5-23C.H 14398.1
generator. A line from the gas geLxerator housing to the sequencevalve vent port on the MOV provides a vent to equalize pressure
in the housing to prevent premature opening of the fuel control
poppet. A position indicator assembly consisting of switches and
a potentiometer provides valve position signals through DDAS to
the S-IVB engine panel. The position switches provide signals to
the gas generator valves OPENED and CLOSED indicators during
checkout and prelaunch operations. During flight, the signals from
the switches and potentiometer are telemetered to ground stations
and recorded. During engine start sequence, opening of the gas
generator control valve is monitored on the S-IVB engine panel
as is event measurement (VK 117). The potentiometer indicates the
control valve position from the closed position through the fully
open position (0 to 100 percent). See J-2 Engine Measuring. At
engine cutoff, the fast fill valve vents the pneumatic pressure fromthe control valve, and a spring returns the valve to the closed
position.
Gas Generator Injector Assmebly. The gas generator injector assem-
bly consists of a circular metal plate containing a normally closed,
spring-loaded oxidizer poppet valve and injector, centered within
a fuel injector ring. The purpose of the gas generator injector
assembly is to distribute propellants into the gas generator combus-
tor. The injector is welded to the gas generator combustor, and
the oxidizer poppet and injector is threaded into the gas generatorinjector assembly. During operation, fuel enters the injector assem-
bly fuel inlet, fills a manifold in the top of the combustor, and
flows through drilled passages in the fuel injector ring. Oxidizer
pressure displaces the oxidizer poppet valve and allows oxidizer
flow through the injector to impinge on the fuel flowing through
the fuel injector ring.
Gas Generator Combustor. The gas generator combustor is a cylin-
drical chamber in which the propellants are mixed and burned.
Two spark plugs initiate combustion. The inlet port mates with
the gas generator injector assembly and the outlet port and shortduct section is welded to the fuel turbine manifold. Propellants
entering the combustor are ignited by the spark plugs; combustion
hot gases pass through the combustor outlet into the fuel turbinemanifold.
Exhaust Ducting. The exhaust ducting and turbine exhaust hoods
are welded sheet metal construction. Dual (Naflex) seals are used
in flanges at all component connections. The ducting conducts the
fuel turbopump turbine exhaust gases to the lox turbopump turbine
and subsequently through the heat exchanger and into the thrustchamber exhaust manifold. A second duct from the fuel turbopump
turbine directs exhaust gases through the oxidizer turbine bypassvalve into the thrust chamber exhaust manifold.
Heat Exchanger. The shell-assembly heat exchanger consists of a
duct, bellows, flanges, and coils. It mounts in the exhaust ductbetween the oxidizer turbine exhaust and the thrust chamber ex-
haust manifold. During flight, cold helium flows through one of
5-24
the four coils, which are heated by the flow of exhaust gases through
the exhaust duct, and expands then returns to the lox tank as ullage
pressurant. The remaining three coils are blanked-off.
Oxidizer Turbine Bypass Valve. The oxidizer turbine bypass valve
is a normally open, spring-loaded gate valve mounted in the oxi-
dizer bypass duct. The valve gate is equipped with a nozzle whose
size is determined during engine calibration. The purpose of the
valve is to prevent an overspeed condition of the oxidizer tur-
bopump and to act as a calibration device for the turbopump
performance balance. When the fuel turbopump turbine starts to
spin, the exhaust gas in the turbine exhaust duct passes through
a duct to the oxidizer turbopump turbine. A percentage of the
gas volume bypasses the oxidizer turbine through the open oxidizer
turbine bypass valve and vents through the thrust chamber. During
engine transit into mainstage, pneumatic pressure, directed to theclosing port of the oxidizer turbine bypass valve, closes the valve
to divert the turbine exhaust gases, except for a volume of gas
which passes through the valve gate nozzle, through the oxidizer
turbopump turbine. During engine shutdown, the ignition phasecontrol valve deenergizes and vents the closing pressure from the
oxidizer turbine bypass valve. The normally open port of the
mainstage control valve supplies pressure to the valve opening
control port to assist the spring in opening the valve. A poten-
tiometer and position switches monitor the bypass valve gate posi-
tion and provide signals through the DDAS to the ESE. The position
switches provide inputs to illuminate the LOX TURBINE BYPASSCLOSED, LOX TURBINE BYPASS OPENED, and LOX TUR-
BINE BYPASS OPEN indicators on the S-IVB engine panel.
During flight these indications are telemetered back to the ESE
as events when the bypass valve closes and opens during the J-2
engine start and cutoff operations. These signals are also monitoredon the events panel. The potentiometer provides an analog signal
of the valves' position from fully opened to fully closed (0 to 100
percent). See J-2 Measuring for additional information.
Control System.
The control system includes the pneumatic control package, whichcontrols helium flow to the various valves in the engine system;
the electrical control package, which controls electrical signals in
the system, and electrical harnesses.
Pneumatic Control Package. The pneumatic control package is a
combination of two regulators, two relief valves, an actuator assem-
bly, a filter unit, and four solenoid valves. The purpose of the
pneumatic control package is to control the flow of helium to the
engine control system and to supply opening and closing control
pressure for all pneumatically operated valves. The control package
also supplies helium for the oxidizer dome, gas generator oxidizer
injector, and oxidizer intermediate seal purges. Helium from the
integral start tank flows into the pneumatic control package through
a filter and a roughing regulator, which reduces the helium pressure
to 450 psi. A spring-loaded, ball-type relief valve that relieves at
3800 psi and reseats at 3500 psi, will bleed off excessive pressure
at the control package inlet. At engine start command, the helium
control valve opens and loads the primary regulator dome with
450 psi. The control regulator samples the output of the primary
regulator and controls primary regulator positions to maintain
system pneumatic control pressure at 400 psi (approximately).
Helium exits the pneumatic control package through a check valve
to the mainstage control valve, the ignition phase control valve,
the pressure accumulator, and to the engine purge lines. Maximum
flowrate through the regulators is 1300 scfm. A low pressure relief
valve prevents over-pressurization of the control system. Cracking
pressure is 497 psig, and reseat pressure is 420 psig. The heliumcontrol and helium tank emergency vent valves are three-way,
electrically operated solenoid valves. The mainstage, ignition phase,
andSTDVcontrolvalvesarefour-way,electricallyoperatedsolen-oidvalveswithopeningandclosingfunctionsarrangedsothatoneisventingwhiletheotherispressurizing.TheSTDVcontrolvalveisnotpartofthepneumaticcontrolpackagebutismountedonandcontrolsthestarttankdischargevalve.TheignitionphasecontrolvalveandtheheliumcontrolvalveareenergizedbytheJ-2enginestartcommand.AnSTDVtimercontrolstheSTDVcontrolvalve,andthemainstagecontrolvalveisenergizedbyanignitionphasetimerduringtheenginestartsequence.Duringcheckoutandprelaunchoperations,commandstoenergizethesolenoidsaremonitoredontheS-IVBenginepanelthroughtheDDAS.ComponentstestswitchesontheS-IVBenginepanelpermitmanualactuationofallfoursolenoidsasnecessaryduringcheckout.Inflight,commandsthatenergizethevalvesaretelemeteredbacktogroundreceivingstationsandrecordedaseventmeasurementsinadditiontobeingdisplayedontheenginepanel.Fast Shutdown Valve. The fast shutdown valve vents the gas genera-
tor control valve permitting it to close and vent the turbine bypass
valve permitting it to open during engine shutdown. Pneumatic
pressure from the normally open port of the ignition phase control
valve opens the normally closed two-position poppet valve. Thevalve is mounted on the main fuel duct at the fuel turbopump
outlet.
Purge Control Valve. The purge control valve is used to control
the thrust chamber oxidizer dome and gas generator purges. The
valve is actuated by pneumatic pressure entering the control port,
acting upon the piston to overcome the force of the piston spring,
and moving the piston to open position. When the valve is notactuated, the force of the spring maintains the valve in closed
position and the outlet side of the valve is vented through the
vent port.
Pneumatic Accumulator. The pneumatic accumulator is an integral
part of the primary instrumentation package, which is enclosedwithin an outer shell and cover. The volume between the primary
instrumentation package and the outer shell and cover serves as
the pneumatic accumulator. The purpose of the pneumatic ac-cumulator is to provide the necessary gas volume for the safe
operation and shutdown of the engine in the event of a loss of
high pressure pneumatic supply.
Electrical Control System. The electrical control system consists of
a sequence controller to properly sequence engine start and cutoff,
and a spark ignition system to establish ignition in the gas generator
and in the augmented spark igniter chamber. The purpose of the
electrical control system is to control engine operation by means
of electrical signals and to supply power to establish ignition. The
electrical control package (figure 5-24) is a sealed, dome-shaped,
pressurized control assembly, containing spark exciters and se-
quence controller circuitry. At engine start or cutoff, the sequence
controller performs the necessary sequencing and timing functions
required to properly operate the engine system. The electrical
control package circuitry will automatically reset for restart capabil-
ity. The sequence controller, mounted in the electrical control
package, is composed of solid-state switching elements, which per-
form the necessary logic functions to properly sequence, time, and
monitor the engine system. The system is completely self-contained
and requires only dc power and external engine start and cutoff
signals for operation. Additional signals are provided to the stageto allow monitoring of the engine condition at significant points
of engine operation. This system also has the capability to actuate
individual components through properly designed checkout equip-ment.
Thrust OK Pressure Switches. Mainstage thrust OK pressure
switches cconsist essentially of an inlet port, a checkout port, 2
diaphragms, toggle blades, a toggl e spring, a housing, and anelectrical switch and connector. Pressure entering the inlet port
Section V S-IVB Stage
ELECTRICALCONTROLPACKAGE/_ THRUST CHAMBER
EXCITER CABLE
RECEPTACLEBLEEDER PLUG _. --v0.,CLECABLE
Jl /- ,_ Jr" RECEPTACLE
%\ _ I/ .fl, _I_ESSU.lZINGTHRUST CHAMBER \\ \ _L_ __ _" I_,'_ Y "\\ \ "_ll"_'r_ _ -/_l _ FITTING
EXCITER CABLE _ _"_3 _'_,,,_!,_ I
RECEPTACLE___ ._'_;_.__,,_ pJ
ENGINE COMPONENT _ _/i$....___'_
CABLE RECEPTACLE _ _ _ GAS GENERATOREXCITER CABLE
RECEPTACLE
GSE CABLE Jj
C-H 14403
Figure 5-24
acts upon a diaphragm linked to the electrical switch through the
toggle spring and toggle blades. As engine oxidizer injector pressureincreases, electrical continuity is switched from normally closed
contact to normally open contact and an electrical circuit is com-
pleted for producing a mainstage OK signal. If oxidizer injector
pressure deteriorates, the pressure switch deactuates, breaking the
contact, interrupting the mainstage OK signal, and re-establishing
a mainstage OK depressurized signal. Proper operation of the switch
may be verified by applying pressure to the checkout port which
is independent of the inlet pressure port. The two switches are
mounted opposite each other on the injector at the lox inlet. A
sensing line attached to each pressure switch and the customer
connect panel (engine/stage interface) permits remote checkout.
Signals from the pressure switches are monitored on the S-IVB
engine panel. During flight, actuation and deactuation signals fromthe switches are telemetered to ground receiving stations as event
measurements. The MAINSTAGE SW NO. 1 PRESS and MAIN-STAGE SW NO. 2"PRESS indicators on the engine panel and
events panel illuminate when lox injector pressure is above 500
± 330 psi. The MAINSTAGE NO. 1 DEPRESSURIZED andMAINSTAGE NO. 2 DEPRESSURIZED indicators illuminate
when lox injector pressure is below the 375-psig minimum actuating
pressure of the thrust OK pressure switches.
Flexible Armored Harness. The flexible armored harness consists
of Teflon-insulated wires that terminate in modified RD (MS R
series) connectors. The conductor wires are wrapped in a layer
of Mylar tape, sleeved inn a silicone rubber tube, and sheathed
in two layers of nickel-plated copper wire braid. The silicone rubber
tube is for thermal protection and the wire braid is for protection
against abrasion and radio frequency interference. Mylar tape is
used only to facilitate installation of the silicone rubber tube. After
installing the wire braid sheaths, exposed ends of braid are soldered
to unify the strands. The wires are soldered or brazed to pins ofthe connectors sealed in the connector housing by neoprene rubber
grommets. To smooth the contour of the harness, potting compoundis applied to each Y-joint and connector prior to installing the
wire braid sheaths. A compound of polyurethane is overmoulded
at the Y-joints and connectors to secure and cover the braid pigtails.
When the harness is installed on the engine, a thermoprotectiveboot is installed over each connector.
Start System.
The start system is comprised of an integral helium and hydrogen
5-25
SectionVS-IVB Stage
start tank, which contains hydrogen (GH 2) and helium gases for
starting and operating the engine; a start tank discharge valve,
which contains the GH 2 in the tank until engine start; a helium
fill-check valve and tank support and fill valve package, whichsupply gases to the start tank; and a vent and relief valve, which
relieves pressure or drains the hydrogen start tank.
Integral Helium and Hydrogen Start Tank. The integral tank consists
of a 4.2 ft _ sphere for GH 2 and a 0.58 ft 3 sphere inside for helium.
GN 2 stored under 1250 psig and 200 ° F provides the energy source
for starting the engine while helium stored under 3000 psig supplies
the pneumatic requirements for engine control system operation.
Since a restart capability is not required for the SL-2 mission, a
start tank inflight refill line has been blanked off. Insulation cover-
ing the start tank prevents excessive internal pressure buildup fromeffects of external temperatures.
During prelaunch operations the integral start tank is purged with
helium, and then pre-chilled, and filled with cold GH z and helium.
Replenish supply is maintained to the helium tank and to the GH 2
start tank until T-3 see. GH z enters the start tank through the
tank support and fill package mounted on the start tank. GH 2
from the ground source, or from the engines having repressurization
capability, flows through poppet check valves into the tank. The
recharge fill port contains a filter to remove contaminants from
GH 2 entering the start tank from the thrust chamber.
Helium enters the helium tank through the helium cover and fill
check valve. The cover contains a tank support for mounting the
start tank, an outlet line to the pneumatic control package, and
a mount for temperature transducer VXC7-401. The poppet-type
fill check valve prevents pressure loss at umbilical disconnect.
Start Tank Discharge Valve. A pneumatically controlled,spring-closed STDV contains the GH 2 in the start tank until needed
for engine start. A control valve directs pneumatic pressure from
the main-fuel-valve sequence valve to open and close the STDV.
Until the start tank discharge delay timer expires during the engine
start sequence, the STDV control valve maintains pressure to the
closing side of the STDV actuator. When the timer expires, the
control valve vents the closing actuator and pressurizes the opening
actuator. GH z then flows through the STDV to the fuel turbopump
turbine. A gate-type check valve on the STDV outlet prevents
gas-generator combustion products from entering the STDV and
contaminating the STDV poppet. Open and closed position switches
and a potentiometer monitor the STDV position. During checkoutoperations, the STDV position (START TANK DISCHARGE
OPEN or START TAANK DISCHARGE CLOSED) is monitored
on the S-IVB engine panel. During flight the STDV positions aretelemetered back to KSC and recorded. The STDV control valve
opening command and the STDV OPEN signals are monitored
in real time on the S-IVB engine panel. See J-2 Engine Monitoring.
Start Tank Vent and Relief Valve. The STDV vent and relief valve
controls GH z flow from the start tank. During servicing operations,
control pressure from the S-IVB pneumatic control system opens
the vent and relief valve to permit GH 2 flow through the start
tank for chilling. The GH 2 exits the start tank through the vent
and relief valve and flows through umbilical lines to the facility
burn pond. Removal of the control pressure permits the STDV
to fill with GH z. During flight the valve relieves excessive pressure
in the start tank. Cracking pressure is 1395 psi minimum and reseat
pressure is 1335 psi.
Start Tank Emergency Vent Valve. The start tank emergency vent
valve provides a redundant means of venting pressurized gas from
the start tank. The valve is actuated, in an emergency, from a
ground source. The emergency vent valve is a solenoid-operated,
two-way, spring-loaded to the closed position, poppet-type valve
mounted on the start tank support-and-fill valve. The valve spring
holds the poppet on the seat, against start tank pressure, when
5-26
PRIMARY FLIGHT
INSTRUMENTATION PACKAGE
D9-401
OXIDIZER PUMP
DISCHARGE r-.- VXD 17-401
PRESSURE'- 7 /START TANK
/ /PRESSURE
DI-401 _"-- BLEEDER / /(TF1)THRUST CHAMBER
\ ,F {_ _a _ GAS GENERATORPRESSURE _ \PLUG / I_ _-D10-40!
_ _,-K/_'_"_ _1 CHAMBER PRESSURE
08--401
FUEL PUMP _ _ 1_7J_'_r'_'_ TM VXDI9-401
.EL,UMTANK
J Jl01JlO0 ,_ _"-- ;NEUMATIC
PRE S R,.NGFITTING
C-H 14402
Figure 5-:)5
the solenoid is deenergized. When the solenoid is energized, the
armature moves to overcome the spring pressure, allowing theflexure to unseat the poppet and vent start tank pressure throughthe valve outlet. The outlet is connected to a line that is teed to
the fuel turbopump primary seal drain line.
J-2 Engine Measuring.
The J-2 engine measuring system, consisting of a primary instru-
mentation package (figure 5-25), an auxiliary instrumentation
package (figure 5-26), and transducers, monitor 36 conditions in
AUXILIARY FLIGH
INSTRUMENT PACKAGED4-401
MAIN FUEL
INJECTOR
PRESSURE-_
MAIN OXIDIZER \
D18-401
ENGINE REGULATOR
OUT PRESSURE_
DUMMy
TRANSDUCER
(3 PLACES) _-/
/--CX,%,%rUR.NE?OX,DIZERTU.,NE/OUTPRESSURE\tN PRESSURE / /---PROPELLANT UTILIZATION
\ / / DVALVEO UTLET PRESSURE
JIS2
ENCLOSURE
NOTE: TRANSDUCERS WITHOUT
CALLOUTS NOT USED ON
S-IVB- 206
CH-14401-I
Figure 5-26
6 parameters and 27 event signals generated by valve positioning
and sequencer commands. The primary instrumentation package
provides instrumentation for monitoring critical engine parameters
during static testing and vehicle flight. The auxiliary instrumenta-
tion package monitors non-critical engine parameters. System de-
sign allows for substitution or deletion of auxiliary package func-
tions without affecting primary instrumentation function. Temper-
ature sensors located in the area of temperature samplings, pressure
transducers located in the instrumentation packages with sensing
lines to the sampling points, flowmeters located in the lines, and
position indicator portions of the valve assemblies acquire the
measurement data and supply it to the S-IVB stage telemetry
system. The S-IVB stage telemetry system telemeters the measure-
ment information to ground receiving stations for recording.
Some measurements are monitored in real time by indicators and
gages on panels, others are displayed on analog recorders. See figure
5-27 for J-2 engine measurement summary. Measurements are
taken in the following parameters: temperature, C; pressure, D;
flow, F: position, G; events, K; voltage, M; and RPM, T. Two-
thousand-ohm potentiometers and/or position switches provide
position indicator signals for the main oxidizer, main fuel, gas
generator control, oxidizer turbine bypass, start tank discharge,
mixture ratio control, augmented spark igniter oxidizer, and propel-
lant bleed valves. Both turbopumps are equipped with magnetic
pickups to measure turbopump speed and to provide a turbine
overspeed cutoff signal for static testing. The magnetic pickups
are utilized to provide turbopump speed measurements for the
instrumentation system. The output of the magnetic transducers
is designed for generation of a 1- to 3-volt pulse suitable for directtelemetry. The fuel turbopump rotor is fabricated from K-monel,
which does not exhibit magnetic qualities until chilled to -300 °
F. Therefore, checkout of the measurement by spinning the tur-
bopump is not feasible at ambient temperatures. Electrical checkout
can be accomplished, however, by applying a voltage to the check-
out coil and inducing a voltage in the signal coil. This check may
be made at either ambient or cryogenic temperatures. Flowmeters
provided within the high pressure propellant discharge ducts mea-sure main fuel and main oxidizer flowrates. The basic element
of the flowmeter is a helical-vaned rotor, which is turned by propel-
lant flow to measure flow velocity. The flow diameter is closely
controlled to permit accurate determination of the volumetric
flowrate. Within the fuel flowmeter is a four-vane rotor that pro-
duces four electrical impulses per revolution and turns approxi-
mately 3600 rpm at nominal flow. The oxidizer flowmeter includes
a six-vane rotor producing six electrical impulses per revolution
and turns approximately 2400 rpm at nominal flow. The output
of the magnetic transducers is designed for generation of a 1- to
3-volt pulse suitable for direct telemetry. Electrical checkout of
the flowmeter can be accomplished by supplying a voltage to the
checkout coil and inducing a signal in the measurement coil. The
12 event (K) measurements are discrete measurements that indicate
when the various engine functions occur. These measurements are
telemetered to the ground receiving stations and recorded. See figure
5-28. The measurements with a 'V' prefix are monitored in real-time
on the LCC-SIVB engine panel. Mission control center monitors
flight control measurements in real-time. These displays assist the
flight controller in making decisions affecting the mission and safetyof the crew.
Acceptance Firing.
The S-IVB-206 stage was installed in the Sacramento test center
beta complex test stand III on June 30, 1966 for acceptance testing.
Prefiring checks on the propulsion system included a manual stage
and GSE control checkout, a system leak check, an automatic
checkout, and a final leak check procedure. Acceptance firing of
the stage was conducted on August 19, 1966. After the propulsion
Section V S-IVB Stage
NAME
At2-403 ACCELERATION, GIMBAL BLOCK
C1-401 TEMF, FUEL TURBINE INLET
C2-401 TEMP, OXID TURBINE INLET
VXC6-401 TEMP, GH 2 START $OTTLE
VXC7-401 TEMP, ENGINE CONTL He
XCll-401 TEMP, ELECT CONTL PKG
XC12-401 TEMP, OG FUEL BLD VLV
C 133-401 TEMP, OXID PUMP DISCHARGE
C134-401 TEMP, FUEL PUMP DISCHARGE
C197-401 TEMP, PR1 INSTR PKG
C198-401 TEMP, AUX INSTR PKG
VXC199-401 TEMP, THRUST CHAMBER JACKET
XC200-401 TEMP, FUEL INJECTION
C21S-401 TEMP, OXID TURBtNE OUTLET
D1-401 PRESS, THRUST CHAMBER
D4-401 PRESS, MAIN FUEL INJECTOR
D5-401 PRESS, MAIN OXID INJECTOR
D7-401 PRESS, OXID TURB INLET
D8-401 PRESS, FUEL PUMP DISCHARGE
DR-401 PRESS, OXID PUMP DISCHARGE
D10-401 PRESS, GG CHAMBER
*VXD17-401 PRESS, GH 2 START BOTTLE
*D18-401 PRESS, ENG REG OUTLET
*VXD19-401 PRESS, ENG CONTL He SPHERE
D57-401 PRESS, PU VALVE OUTLET
D86-401 PRESS, OXID TURBINE OUTLET
XD241-401 PRESS, GH 2 START- BOTTLE, BACKUP
XD242-401 PRESS, ENG CONTL He SPHERE, BACKUP
D266-401 PRESS, THRUST CHAMBER OSClLL
F1-401 FLOWRAT E, OXlD
F2-401 FLOWRAT E, FUEL
*G3-401 POSITION, MAIN OXID VALVE
*G4--401 POSITION, MAIN FUEL VALVE
G5-401 POSITION, GAS GEN VALVE
*G8-401 POSITION, OXID TURBINE BYPASS VALVE
G9-401 POSITION, GH 2 START VALVE
*VXGIT-401 POSITION, MIXTURE RATIO CONTROL VALVE
M6-401 VOL'i:AGE, ENGINE CONTROL BUS
M7-401 VOLTAGE, ENGINE IGNITION BUS
T1-401 SPEED, OXID PUMP
T2-401 SPEED, FUEL PUMP
MEASUREMENT NUMBER PREFIXES INDICATE THE FOLLOWING:
*FLtGHT CONTROL; V-ESE DISPLAY; X-AUXILIARY DISPLAY
RANGE
0TOSG
460 TO 2260°R
460 TO 1660°R
110 TO 560°R
110 TO 560°R
160 TO 660OR
35 TO 85OR
160 TO 210°R
35 TO 60°R
160 TO 660°R
160 TO 660°R
35 TO 560°R
35 TO 560°R
440 TO 1460°R
0 TO 1000 PSIA
0TO 1000 PSIA
0 TO 1000 PSIA
0 TO 200 PSIA
0TO 1500 PSIA
0 TO 1500 PSIA
0 TO 1000 PSIA
0 TO 1500 PSIA
0 TO 750 PSIA
0 TO 3500 PSIA
0 TO 500 PSIA
0TO 100 PSIA
0 TO 1500 PSIA
0 TO 3500 PSIA
-5 TO +S PSIA
0 TO 3000 GPM
0 TO 9000 GPM
0 TO 100%
0 TO 100%
0 TO 100%
0 TO 100%
0 TO 100%
0 TO 65 DEG
0 TO 30V
0 TO 30V
0 TO 12 K RPM
0 TO 30 K RPM
CH.14399.1
Figure 5-27
5-27
Section V S-IVB Stage
NUMBER
VK5-401
VK6-401
K7-401
VXK8-401
VK10-401
VK11-401
*K12-401
*VXK13-401
*VK14-401
K20-401
VK95-401
VK96-401
Kl16-401
VKl17-401
VKl18-401
Kl19-401
VK120-401
K121-401
VK122-401
K123-401
VK124-401
VK125-401
*K126-401
*KT27-401
*VK 1E7-401
K158-401
K159-401
NAME
MAINSTAGE CONTL SOL ON
IGNITION PHASE CONTL SOL ON
HELIUM CONTL SOL ON
IGNITION DETECTED
TC SPARK SYS ON
GG SPARK SYS ON
ENGINE READY SIGNAL
CUTOFF SIGNAL
MAINSTAGE OK PRESS _ 1
AS1 LOX VALVE OPEN
TC INJECTORTEMP OK
STDV CONTL SOL ON
GG CONTL VALVE CLOSED
GG CONTL VALVE OPEN
MAIN FUEL VALVE OPEN
MAIN FUEL VALVE CLOSED
MAIN OXlD VALVE OPEN
MAIN OXlD VALVE CLOSED
STDV OPEN
STDV CLOSED
OXlD TURBINE BYPASS VALVE OPEN
OXlD TURBINE BYPASS VALVE CLOSED
OXID BLEED VALVE CLOSED
FUEL BLEED VALVE CLOSED
MAINSTAGE OK PRESS SW 2
MAINSTAGE OK PRESS SW I DEPRESSURIZED
MAINSTAGE OK PRESS SW 2 DEPRESSURIZED
MEASUREMENT NUM_R PREFIXES INDICATE THE FOLLOWING:
*FLIGHT CONTROL; V-ESE DISPLAY; X-AUXILIARY DISPLAy
CH-14400.1
Figure 5-28
system achieved 436.1 sec of mainstage operation, a computer
controlled program issued the cutoffcommand. Major engine events
of the acceptance test firing included:
a. Normal engine start sequence.
b. Propellant utilization system activation 6 sec after engine startcommand.
c. Engine side load restrainer links were released approximately25 sec after mainstage control.
d. An automatically controlled gimbal program was initiated afterrestrainer link release.
e. After 300 sec of mainstage control, the propellant utilization
valve repositioned from the lox-rich approximate mixture ratio of
5.5 to 1.0 to the nominal position for reference mixture ratio of
4.7 to 1.0 for duration of firing.
f. After 436.1 sec of mainstage operation, automatic cutoff was
initiated at 2216 Ibm of lox, or 596 Ibm of LH 2. Depletion sensors
armed for 3-percent residual mass served as backup for the cutoff.
g. Post-firing inspection revealed that metal chips were present
in the turbine exhaust system and a damaged turbine seal was
discovered. The lox turbopump was replaced and a 70 sec verifi-
cation firing at 5.5 EMR was conducted on September 14, 1966.
Postfiring propulsion system checks included test equipment remO-
val, leak check of the system, and system automatic checks. Inspec-
tion checks reverified engine alignment and structural integrity after
the static acceptance firing test.
PROPELLANT SYSTEMS.
The S-IVB propellant tanks were sized to accommodate the Saturn
V mission requirements. The propellant masses loaded provide
hydrogen and oxygen for J-2 engine operation, boiloff, LH 2 tank
pressurization, and usable and unusable residuals. The propellant
tank fill systems were designed to minimize countdown and pre-
launch time and to be compatible with the loading times scheduled
for the other stages of the Saturn V vehicle. The fill systems were
sized to flow 1000 gpm of liquid oxygen and 3000 gpm of liquid
hydrogen. Initial fill rates are slower to accomplish tank chilldown
and to prevent ullage pressure collapse. Final fill rates are also
slower to provide greater precision in obtaining the desired 100%load.
The propellant tank vent systems were designed to protect the tank
structure under all conditions of propellant tank loading, stage
powered flight and orbital venting. During loading, tank pressures
are maintained well below the normal tank prepressurization levels.
The negative pressure differential across the common bulkhead
between the fuel and oxidizer tanks is the limiting factor. This
differential pressure is never allowed to exceed 3.0 psi for normal
controlled operation. Vent outlets are so located and directed as
to prevent disturbing moments on the stage during venting in flight.
Propellant Characteristics.
Figure 5-29 lists the physical and chemical properties of LH 2, and
figure 5-30 shows the LH 2 vapor pressure curve. See the S-IB
stage lox system description in Section IV for lox characteristics.
Fuel Tank.
The LH 2 is stored in an insulated tank with a capacity of approxi-
mately 44,300 Ibm at a temperature of-423 ° F. An ullage volume
of approximately 400 ft _ is maintained at this load level; however,
for Saturn IB earth orbit missions a full load of LH 2 is not carried.
The approximate LH 2 load for S-IVB-206 is 38,000 Ibm, which
includes quantities required for programmed mixture ratio opera-
tion (5.5 to 1 for approximately 325 see) and unusable propellants.
The LH 2 tank is pressurized between 28 and 31 psia. Pressure
is provided by a ground supply of helium for prepressure and is
maintained by a hydrogen bleed from the engine during burn.
Tank venting and relief is accomplished through parallel valves
installed in a top-mounted vent system which exits through nonpro-pulsive vent outlets in the forward skirt. An anti-vortex screen is
installed over the engine feed duct inlet.
Fuel Loading.
Chilldown. The automatic fuel loading sequence is started 4 hr
and 19 min before liftoff by manually depressing the fill pushbutton on the I_,H 2 control panel in the storage facility. The chill-
down sequence begins and the following five operations occursimultaneously.
a. The vehicle LH 2 vent-and-relief valve opens (figures 5-31 and
5-32). W;_h ',he directional control valve now in the ground position,
the vent gases flow through the umbilical to the facility burn pond.
b. The facility line chilldown valve, fill valve (on LH 2 control sled),
5-28
COMMON NAME:
LIQUID HYDROGEN (LH2)
CHEMICAL FORMULA:
H2
MOLECULAR WEIGHT:
2.016
PHYSICAL PROPERTIES:
FREEZING POINT ..................... -435°E
BOILING POINT ...................... -422.9°F
CRITICAL TEMPERATURE ............. -399.96°F
CRITICAL PRESSURE ................... -188,16 PSIA
LIQUID DENSITY ...................... 0.594 LBS/GAL
APPEARANCE ........................... CLEAR LIQUID
ODOR .................................... NONE
LIQUID TO GAS RATIO .............. 1:780
FLAMMABILITY LIMITS ............... 4°1oto 74.5%BY VOLUME
IN AIR
CHEMICAL PROPERTIES:
STABLE AGAINST MECHANICAL SHOCK IN PURE FORM.
IMPACT SENSITIVE IF CONTAMINATED WITH LIQUID AIR OR OXYGEN ICE.
HIGHLY FLAMMABLE WHEN MIXED WITH AIR, LOW IGNITION TEMPERATURE,
_061 - 1065°F, CREATES EXTREME FIRE HAZARD.
HELIUM ONLY MATERIAL THAT WILL NOT TURN TO A SOLID WHEN IMMERSED
IN LH 2.
C-H 14541
Figure 5-29
and LH 2 debris valve (on service arm 6) open; and the vehicle
fill and drain valve opens establishing a flow path from the storage
facility LH 2 tank to the vehicle fuel tank.
c. The gas heat exchanger inlet valve (on the umbilical tower) opens
and internal level sensors become active filling the heat exchanger
with LH z to a controlled level. See Lox Pressurant Loading for
gas heat exchanger function.
d. A 4-min facility timer starts.
e. The helium nozzle purge starts. (Purge of the LH 2 fill and draindisconnect).
f. At the end of four min, the facility storage tank pressurization
valve opens allowing LH 2 to flow through a vaporizer producing
GH 2 that flows into the storage tank ullage space providing pres-
sure to force out the LH z for vehicle loading.
As pressure builds in the LH 2 storage tank, and at the end of
four min. LH z begins'flowing through the transfer line increasing
to 500 gpm. This flow chills the transfer system including the vehicle
LH 2 tank.
Initial Fill. When the facility 4-min timer expires, the facility fill
valve partially closes, the line fill valve opens, and the line chilldownvalve closes starting the initial fill sequence. LH 2 flows into the
vehicle tank at 500 gpm until the PU (propellant utilization) mass
probe signals the LH z tanking computer that LH 2 has reached
the 5-percent level. This requires about 7 min.
Main Fill. When LH 2 reaches the 5-percent level, the tanking
computer commands the facility fill valve to fully open increasingthe flow rate to 3000 gpm. When either the PU mass probe or
the fast fill sensor indicate the LH 2 has reached the 96-percent
level (after approximately 24 min). the tanking computer starts
the slow fill sequence.
Slow FilL The facility fill valve partially closes and the replenish
Section V S-IVB Stage
valve (on LH 2 control sled) opens establishing a flow of 500lbm/min until the LH 2 level reaches 100-percent. This requires
approximately 9 min after which the tanking computer starts the
replenish sequence.
Replenish. The facility fill valve closes and LH 2 flow through the
replenish valve maintains the tank level at 100 percent. Should
boiloff exceed the replenish valve capacity and drop the level below
90 percent, the tanking computer will partially open the facility
fill valve until the level again reaches 100 percent.
Umbilical Drain. Replenish ends and umbilical drain begins at start
of automatic sequence. Then the vehicle LH 2 fill and drain valve
closes, the umbilical line drain valve opens, and helium purges
the umbilical line. At liftoff the service arm debris valve closes.
Fuel Tank Pressurization.
Prepressurization. The prepressurize-for-launch command l min 37
sec before liftoff starts 800-psig ground helium flow through the
stage pressurization system (figures 5-33 and 5-34) into the tankullage space to assure the necessary head at the engine pump at
ignition. When the tank ullage reaches 31 psia, a pressure switchactuates, sending a signal to close the ground prepressurization
supply valve. If the pressure drops to 28 psia, the switch willdeactuate opening the supply valve replenishing the ullage gas.
After liftoff no pressurization supply is required until engine igni-tion after which the flight pressurization system is enabled. Over-
pressurization during any period, though not expected, would re-
lieve through mechanical operation of the vent and relief valve
or latching vent and relief valve in the 31 to 34 psia band. During
flight with the directional control valve in the flight position, vent
gases pass through two 4-in. ducts that exhaust through the skirt
directly opposite each other for total thrust cancellation.
Flight Pressurization. GH 2 bled from the J-2 engine flows to the
pressurization control module, which has three orificed flow paths
and into the ullage space of the fuel tank. One path is always
open, but normally-open solenoid valves control the other two.A switch selector command (channel 68) 2.6 sec after J-2 ignition
closes one solenoid valve and enables the circuit so the flight control
HYDROGEN rAPOR PRESSURE
50
Z
c¢
Z
160-
140-
t20-
100-
80-
60-
40-
20-
0
-440
• D
TEMP. OF PSIA
.0 0.00 FREEZING POINT
A -422,9 14.70 BOILING POINT(I ATM.)
S o -399 96 188.16 CRITICAL
i / i i _ i i i i i _ i i4,30 420 410 400 390 380
MINUS DEGREES FAHRENHEIT (OF) C-H 14S40
Figure 5-30
5-29
SectionVS-IVBStage
S-IVB FUELSYSTEM
LEGEND
L_
HELIUM IIIIll
ELECTRICAL_
FUEL FILL &
DRAIN DISCONNECT
FUEL | |
DEBRIS VALVE' | |
LH 2 CONTROL FACILITY | |I II I
LH 2 CHILLDOWNNetes
J_ SWITCH SELECTORSEE ELECTRICAL SYSTEM
:> SEE MEASUREMENT TABLE,L
SEE PNEUMATIC CONTROL SYSTEM C_LLDOWNSHUTOFF
OPEN INDICATION
iFUEL BLEED Vt.V CLOSE n
FUEL CHILLDOWN
INVERTER
XD2
IIII
"L
NK
I !
II
Figure 5-31 CH-14528.2
5-30
pressure switch can control the other valve. When the tank pressure
is above 28 psia, the pressure switch actuates holding the valve
closed restricting the GH 2 flow to a single path. When the tank
pressure drops below 28 psia, the switch deactuates allowing the
valve to return to the normally-open position providing two GH 2
flow paths. After 5 min 0.2 sec, another command (channel 69)
disables the control circuits opening both controlled paths and
allowing maximum GH 2 flow during the last portion of J-2 engine
burn (about 2 min 15 sec) when the liquid level reaches a point
that bulk temperature stratification requires increased ullage pres-
sure to meet engine inlet requirements.
LH= Chilldown.
During fuel loading, the 10-in. prevalve remains open (its normal
position) providing a partial chilldown of the feed system. At 10min before liftoff, the LH 2 chilldown pump is started. An interlock
is provided to ensure that the vehicle LH 2 tank is at least 10 percent
full prior to energizing the pump. The prevalve is closed 5 sec
after the pump starts. The LH_ is now routed from the S-IVB
stage tank, through the pump, chilldown shutoff valve, line strainer,
and J-2 engine feed duct, fuel turbopump and fuel bleed valve,
and returned to the LHz tank through the return line and check
valve. The pump delivers 135 gpm with a differential pressure of
6 to 9 psi. This recirculation is required tomaintain a liquid phase
in the turbopump to ensure proper fuel quality for engine start
and to prevent turbine overspeed and possible disintegration at
engine start. The LH 2 chilldown pump operates continuously until
0.6 sec prior to J-2 engine start. At 4.5 sec before J-2 start command,
a switch selector command (channel 83) opens the LH z prevalve.
Since the chilldown pump is still operating, the LH 2 is diverted
from the normal circulation path and flows back through the LH 2
prevalve and into the fuel tank. The backflow of LH 2 clears any
vapor entrapment from the fuel tank outlet area.
Fuel Measurements.
Figure 5-35 lists the fuel system flight measurements and indicatesthe information that is monitored in real time. Measurements
VXD177 and VXD178 are also displayed in analog form on the
command module main display console providing fuel tank pressuredata for the crew.
Section V S-IVB Stage
Main FilL When the level reaches 5 percent (sensed by the PU
lox mass probe), the flowrate is increased to a maximum of 10,494
lb/min. After approximately 22 min, the level reaches 94 percent
and slow fill begins.
Slow Fill. When the level reaches 94 percent, the flowrate is reduced
to a maximum of 3t46 lbm/min. In approximately 1 min 30 sec,
99 percent of flight mass is loaded.
Replenish. When the level reaches 99 percent, the tanking computer
enables the lox replenish flow control valve. The computer controls
the replenish flowrate to match the boiloff rate. The maximum
replenish flowrate is 575 lbm/min. The replenish sequence ends
at the start of the automatic sequence. At that time the lox fill
and drain valve closes, the lox replenish flow control valve closes,and the umbilical line vent valve and the umbilical purge valve
open. The lox debris valve closes at liftoff.
kox Tank PrepressurlzaUon.
The automatic lox tank prepressurization operation begins 2 min47 sec before liftoff. At this automatic command the lox tank
vent-and-relief-valve closes (figures 5-38 and 5-39). The two nor-
mally-closed helium supply valves are opened allowing helium toflow at a rate of 20 lbm/min maximum, at -360 ° F and 2000
psig for 15 sec. The flow is through the aft umbilical into the plenum
chamber, through the normally open J-2 engine heat exchanger
bypass valve, and the pressure regulating orifices, and into the
lox tank. When the lox tank pressure reaches 40 psia, the prepres-
surization and flight control pressure switch controlling the helium
supply shutoff valve actuates closing the valves. If the pressuredrops below 37 psia, the switch deactuates allowing more helium
flow into the tank. The helium ground supply valve remains active
until S-IB ignition command (3 sec before liftoff) when it closesautomatically and the supply line vent valve opens, allowing the
Lox Tank.
The lox is stored in a tank formed by the aft dome and common
bulkhead with a capacity of approximately 194,000 Ibm of lox
allowing for approximately 70 ft _ of ullage. The lox tank is loaded
to capacity for Saturn IB missions. The tank ullage pressure is
maintained between 37.0 and 40.8 psia during boost and engine
operation using GHe. Two parallel valves provide venting and
overpressure relief through the aft skirt. An anti-vortex screen isinstalled over the engine feed duct inlet.
Lox Loading.
The following steps describe the automatic lox loading operationas controlled and monitored in the LCC. See figures 5-36 and 5-37.
Chilldown. When the S-IB lox level reaches 65%, the S-IVB main
line chilldown begins by opening the S-IVB repl.enish valve on
the lox control sled, the S-IVB lox vent valve, the lox debris valve,and the S-IVB lox fill and drain valve. When these valves are
open, main storage tank pressure causes lox to flow slowly throughthe S-IVB main line causing chilldown. When the S-IB lox level
reaches 97%, S-IVB slow fill starts and the tank is filled to 5 percentat a maximum flowrate of 5225 lbm/min. Figure 5-32
5-31
Section V S-IVB Stage
S-IVB FUEL TANK PRESSURIZATION AND VENT SYSTEM
S-18 IGNITION CMD _) ._SEQ pWR TRANSFER CMD (T-28) S-IB CUTOFF CMD
ACTUATION CONTROL
MODULE (3 PLACE._
K 13
+4015
r--SERVICE
TO GAS _ARM NO. 7
BURN POND \
FLIGHT _/%_
NON PROP. VENT
-NON PROP. |
VENT NO. 2J
I
+4DI5
CONTROL
VALVE
RELIEF
VALVE
LEGEND
HELIUM
FUEL TANK ULLAGEIII
ELECTRICAL SIGNAL_
LATCH OPEN
TANK
S-IB CUTOFFm
COMMAND i_'_
_°LH 2 DIR VENT VALVE
.:_. ,NGROUNDPOS,T,ONS-IV8 STAGE /
PRESSURE PANEL _
LrLrLr_2VENT OPEN
I LH2 READY TO LOAD
J LH 2 LOADING J
L SEQUENCE
1 TANKPO._UNCH' I,/-------1÷/--_---__/FIRINGCOMMAND I LH2TANKVENT..4.1/ I / .----_UI--"
Ir_ VALVECLOSEDI " " I ._,/v S-IB CUTOFF COMMAND I
l
I L/ Ieh--_LCj LH2VENTVALVEI ;L// -- BOOST CLOSE
_j (_ PREPRESSURE I LH2 TANK PREPRESSURIZATION-- SUPPLY VENT | VALVE OPEN COMMAND
CLOSE COMMAND(_COMMANDS_'IBIGNITION e_,|j_ _,., SUPPLY LINE VENT OPEN
5-32Figure 5-33
+4DI5
RELIEF
-J-2 ENGINE
+4DI1
D104 PRESSURIZATIONCONTROL MODU1.E
CHECK
UMBILICAL
DISCI
FROM
Notes
[_ SWITCH SELECTOR.SEE ELECTRICAL
SYSTEM
[_SEE MEASUREMENTTABLE
SEE PNEUMATIC
CONTROL SYSTEM
ARM NO. 6
FROM LH 2
STORAGE
GAS HEAT
EXCHAI_
CH.14380.2
helium supply line to become inerted from the stage supply check
valve to the closed helium supply valve.
Lox Pressurant Loading.
Helium is used as the pressurant in the lox tank during flight.
The helium is stored in the vehicle in six spheres manifolded
together in the fuel tank. The filling of these spheres is a ground
controlled three step operation. The first transfer is at a pressure
of 950 -+ 50 psig, ambient temperature, and a flowrate of 5.1
lbm/min maximum. The minimum time required to make this
transfer of 14 Ibm of helium is 4 min. The helium pressure is
regulated as it flows through pneumatic console 432A the gas heat
exchanger, and pneumatic console 433A and then flows through
service arm 6 into the aft umbilical of the stage and into the spheres.
Prior to the lox loading, the second step of helium transfer starts.
The pressure is increased to 1450 ± 50 psig, still at ambient
temperature, and the flowrate is 7.6 lbm/min maximum. During
this step, 7 Ibm of helium is transferred in 2 min 24 sec. The
third step begins after the helium spheres become immersed in
LH z approximately 4 hr before liftoff. The gas heat exchangeris filled with LH 2 and the helium is chilled to -410 ° F as it passes
through. The cold helium pressure is increased to 3100 -q- 100
psig and loaded at a flowrate of 32.5 lbm/min maximum, until
a total mass of helium reaches 246 Ibm. Continuous pressurization
Section V S-IVB Stage
is available to the spheres until 3 sec before launch. The cold helium
dump module contains a solenoid vent valve that will vent the
spheres upon command from the LCC S-IVB stage pressure panel.This module also contains a relief valve that guards against over
pressurization by cracking at 3500 psig. It reseats at 3200 psig.
Lax Tank Flight Pressurization.
During S-IB powered flight, lox ullage makeup cycles are initiated.The cold helium shutoff valves are opened 2.5 sec prior to S-IVB
engine start and remain open throughout burn. Cold helium flows
from the storage spheres through the control module, where it is
filtered and regulated to 385 ( + 28-32) psia, and into the compressed
gas tank plenum. The medium pressure switch senses plenum
pressure and backs up the regulator by closing the two shutoff
valves if plenum pressure reaches 465 psia. When the pressure
decreases to 350 psia. the switch deactuates allowing the shutoff
valves to open again. From the plenum, helium flows to the lox
tank ullage space through three paths: (I) through an orifice and
directly into the tank, (2) through the heat exchanger coil and
an orifice, and (3) through the heat exchanger coil. and the normal-
ly-open heat exchanger bypass valve.
The helium that flows through the heat exchanger is warmed and
expanded. At engine start plus 23.4 sec. the flight control pressure
switch control of the heat exchanger bypass valve is enabled. The
S-IVB FUELTANK PRESSURIZATItlNAND VENT SYSTEMDIAGRAM
NOTES:
J_ SEE S-IVB PNEUMATICCONTROL PRESSURE
SYSTEM DIAGRAM
J_ FROM J-2 ENGINEGH 2
AC.UAT.ONCO "ROL--t-- - mNON-,RO,U'S,V,V...MOOU.E(3P.ACES, Itl
w_,__
__ LH 2 TANK
Figure 5-34
_>
LH 2 VENT
DISCONNECT
tEL TANK
PRESSURIZATION
CONTROL MODULE
3
LH 2 TANK _
PREPRESSURIZATIONDISCONNECT _//
5-33
SectionVS-IVBStage
LIQUID HYDROGEN SYSTEM MEASURMENTS
NUMBER
*XC3-403
C15-410
CS2-408
C 157-404
C 161-424
C254-,109
C255-409
*XD2-403
VD54-410
D104-403
*VXD177-408
*VXD t78-408
*D183-409
*D184-409
"D218-403
*VXF5-404
*K1-410
K3-427
*K17-410
K 19-403
'_Kll 1-404
*Kl12-404
K113-4il
Kl14-411
*K210-410
*K211-410
LI-408
L2-_)8
NAME
TEMP, FUEL PUMP INLET
TEMP, FUELTANK GH 2 INLET
TEMP, FUELTANK POS° I
TEMP, FUEL ClRC PUMP OUTLET
TEMP, LH 2 CIRC RET LINE TANK INLET
TEMP, LH 2 TANK NON-PROP VENT I
TEMP, LH 2 TANK NON-PROP VENT 2
PRESS, FUEL PUMP INLET
PRESS, FUEL TANK INLET
PRESS, LH 2 PRESS MODULE INLET
PRESS, FUEL TANK ULLAGE EDS 1
PRESS, FUEL TANK ULLAGE EDS 2
PRESS, LH 2 TANK NON-PROP VENT 1
PRESS, LH2TANK NON-PROP VENT 2
RANGE
35 TO 47oR
50 TO 300OR
35 TO 47°R
35 TO 50°R
35 TO 50°R
25 TO 260°R
25 TO 260OR
0 TO 60 psla
0 TO 100 psla
0 TO 1000 psla
0 TO 50 psia
0 TO 50 psia
0 TO 50 psia
0 TO 50 psla
PRESS, LH 2 CHILLDOWN PUMP DIFF
FLOWRATE, FUEL CIRCULATION PUMP
EVENT, FUELTANK VENT VALVE CLOSED
EVENT, FUEL FILL VALVE CLOSED
EVENT, FUEL TANK VENT VALVE 1 OPEN
EVENT, FUEL FILL VALVE OPEN
EVENT, FUEL PREVALVE OPEN
EVENT, FUEL PREVALVE CLOSED
EVENT, LH 2 TANK VENT VALVE C CLOSED
EVENT, LH 2 TANK VENT VALVE D CLOSED
EVENT, LH 2 LATCH RELIEFVALVE CLOSED
EVENT, LH 2 LATCH RELIEF VALVE OPEN
LEVEL, FUEL TANK POS I
LEVEL, FUEL TANK POS 2
-30 TO 30 i_;d
0 TO 160 gpm
0 TO 5 vdc
0TO 5 vd¢
MEASUREMENT NUMBER PREFIXES INDICATE THE FOLLOWING:
*FLIGHT CONTROL; V-ESE DISPLAY; X-AUXILIARY DISPLAY CH.20125
Figure 5-35
pressure switch opens and closes the valve as it senses ullage
pressure and regulates the quantity of helium that flows into thetank. A switch selector command (channel 79) one sec after J-2
cutoff opens the circuit and closes the shutoff valves.
Lox Chilldown.
During lox loading, the 10-in. prevalve in the engine feed duct
remains open (its normal position) providing a partial chilldown
of the feed system. Chilldown is required to properly condition
the Iox turbopump and feed duct to meet turbopump inlet require-
ments at engine ignition. A ground command starts the Iox chill-
down pump 10 min before liftoff, and it runs continuously until
0.4 sec before J-2 ignition command. The prevalve closes 5 sec
after the pump starts. The pump circulates lox through the normally
open lox chilldown shutoff valve, the Iox feed duct, the Iox tur-
bopump, the high pressure propellant duct, the Iox bleed valve,and the return check valve back into the tank. The pump delivers
38 gpm at 15 psia. The prevalve opens on a switch selector command(channel 83) 4.5 sec before J-2 start command allowing reverse
Iox flow through the feed duct removing trapped bubbles throughthe duct and anti-vortex screen.
5-34
Lox Measurements.
Figure 5-40 lists the lox system flight measurements and indicatesthe information that is monitored in real time. Measurements
VDI79 and VDI80 are also displayed in analog form on the com-
mand module main display console providing lox tank pressuredata for the crew.
Propellant Utilization.
The PU (propellant utilization) system (figure 5-41) is provided
for mass indications- during loading to within ___+1.0 percent accu-
racy, and inflight propellant mass history to the telemetry system
to an accuracy of _ 1.0 percent of total propellant mass. The PU
electronic assembly, located in the forward skirt, is pressurized and
mounted on an environmental panel. Inputs to the PU electronic
assembly are from mass sensor probes, one located in each propel-lant tank. The probes are cylindrical capacitors, varying in capaci-
tance linearly with the liquid mass in the tanks. Each mass sensor
forms one leg of a servo-balanced capacitance bridge. The output
signal of the bridge drives a servo-motor (15ot positioner) which
positions a potentiometer wiper to yield a signal that is supplied
to the ground loading computer during propellant loading. During
flight, the same signals are telemetered (figure 5-42) to provide
a flight history of propellant masses.
SA-206 S-IVB ORBITAL SAFING.
The propellant tanks will be vented immediately after engine cutoff
to ensure that no stage disturbances oc:cur during CSM separation,
which will occur approximately six rain after S-IVB cutoff. Sub-
sequently, the following sating operations will be accomplished.
a. The lox tank will be vented through the Iox NPV system.
b. The LH z tank will be vented through the LH 2 NPV system.
c. The cold helium bottle residuals will be dumped through the
lox NPV system.
d. The stage pneumatic bottle residuals will be dumped through
the engine pump purge control module.
e. The engine start bottle will not be refilled during engine powered
flight so it will not have to be dumped.
f. The engine control bottle residuals will be dumped through the
engine purge system.
Rapid elimination of these residuals will preclude high tank pres-
sures and result in a completely safe vehicle for orbital coast.
Propellant Tank Orbital Venting.
Orbital venting starts after J-2 engine cutoff. Figures 5-43 and 5-44
illustrate the sequence of events during this period. Switch selectorcommands are sent 0.2 and 0.4 sec after the J-2 cutoff command
to open the LH z and Iox valves, respectively. The Iox tank ventsfor about 3 rain 20 sec while the LH., tank vents for about 5
min 20 sea The LH 2 and lox tank vent systems are non-propulsive.
The non-propulsive vent valves are latched open at 15 rain alter
J-2 engine cutoff to allow continuous tank venting, which precludes
any long term tank pressure buildup.
PNEUMATIC CONTROL SYSTEM.
Pneumatic control of all propulsion system components (lignres
5-45 and 5-46) requiring command actuation was chosen to provide
the capability or rapid response and high force where needed with
minimum weight of hardware required for the power source and
Section V S-IVB Stage
S-IVB LOX SYSTEM
_ _.-LOX TANK
_ VXF4
Nlmmaulmmm mim ammmm mummmmu mmmmmmmmim
m m• !
!Ummma
XC4
CHILLDOWN
INVERTER
.I-2 ENGINE
NOTE
SWITCH SELECTOR.SEE ELECTRICAL SYSTEM
SEE PNEUMATIC CONTROL SYSTEM
SEE MEASUREMENTS TABLE
LEGEND
nmm LOX
li I i HELIUM
ELECTRICAL
Figure 5-36
Jl
CH-14527-2
5-35
Section V S-IVB Stage
S-IVB LOX SYSTEM DIAGRAMOVER FILL
FAST FILL
LOX FILL
AND DRAINFILL AND DRAIN
DISCONNECTj]_ [
I
ACTUATION
CONTROL MODULE
(2 PLACES).
LOX MASS
SENSING
Figure 5-37
control components. A storage pressure of 3000 psi was chosento be compatible with other vehicle requirements and to take
advantage of components previously developed for this pressure.
The pneumatic control system provides supply pressure for stage
pneumatically operated valves including the J-2 engine start tank
vent. Pneumatic power for all the other engine valves is suppliedby the engine pneumatic supply. A pneumatic regulator assem-
bly regulates filtered ambient helium flowing from the ambient
storage sphere at 3000 -+-100 psia and 70 ° F. The module regulates
pressure down to 470_ 12 psig for operation of the LHz directional
control valve, the Iox and LH :_ vent valves, the Iox and LH ._,fill
and drain valves, the J-2 engine G H :_start system vent-relief valve,
the lox and LH e prevalves, the Iox and LH:a chilldown shutoff"
valves. Iox NPV valve, and LH ._,latching relief valve, and for pnr_eof the J-2 engine LH., and Iox turbopunlp turbines and the Iox
chilldown pump housing. Several other components are purged
during preflight operations" for a list of purges _!nd flowrates, see
ligure 5-47.
PREPRESSURIZATION OPERATION.
Prcprcssnrization is accomplished by opening tile helimn dome
supply valve on console 432A and by actuating tile switch oll tile
helium control panel located in tile LCC. See Iigure 5-48 for
pneumatic control system mcusuremcnts und gronnd displays. Tile
storage sphere is pressurized to 1000 psig and the regulator nlain-tains a plenunl cllambcr pressure of 470 psig downslre;|nl of tile
storagc sphcrc. The ground supply helium sonl'ce through pneu-
matic console 432A is maintained prcventing depletion of the
control prcssnrc.
5-36
CHILLDOWN
PUMP
RELIEF
VALVE
J_ _ PURGEVENT
NOTES:.
[_ HELIUM PURGE, SEE S-IVBPNEUMATIC CONTROL SYSTEM DIAGRAM
J_ 470 PSIG FROM S-IVBHELIUM
PNEUMATIC CONTROL SYSTEM
[_ LOX CHILLDOWN RETURN LI NE
FROM J-2 ENGINE
[_ SEE S-IVB LOX TANK PRESSURIZATION &VENT SYSTEM DIAGRAM
FUEL PREVALVE AND CHILLDOWNVALVE CONTROL PRESSURE, SEE
FUEL SYSTEMCH,14713.1
PREFLIGHT PRESSURIZATION OPERATION.
The preflight pressurization begins at approximately 6 hr 37 min
before liftoff. A regulator dome pressure supply solenoid value
in console 432A is opened and helium at 3000 psig flows to the
control helium sphere. The control helium storage sphere is made
of titanium and will withstand u minimum proof pressure of 4800
psig in a temperature range of-40 to 210 ° F without structural
failure. The ground supply pressure is maintained until 3 sec betbre
liftofl'at which time the supply is terminated by ignition comn'tand.
A check vulve maintains subsystem pressure fur inflight operations.
FLIGHT CONTROL.
Two modes of flight control, burn and coast, nlaintuin attitude
control of tile S-IVB/IU/payload vehicle configuration from S-IB/
S-IVB separation until shortly after CSM separation from the
launch vehicle. During the burn mode. attitude and steering adjt, st-
ments are made by hydraulically gimballing tile J-2 engine -----7
deg muxinlnm in tile pitch and yaw phmes. Auxilary propulsion
system (APS) modules on the S-IVB stage ali skirt provide vehicle
roll control during the burn mode by thrusting tangentially to the
vehicle. During the coast nlode, which begins ;It .I-2 engine ctttoff.
the APS modules provide pitch, yaw. and roll attitude control. All
attitude and steering conlnlands originate in tile ILl guidance.
nuvigation, and control syslcm. As conlputations are made tile flight
control computer issues the conlnlands that gimbul the .I-2 engine
;-IVB LOX TANK 'RESSURIZATION AND VENT SYSTEM
i.
+4Dll
VALVE I_
PROPULSIVE
C207
COLD HELIUM
-LOX PRESSURIZATION
CONTROL MODULE
Section V S-IVB Stage
IHELIUM
DISCONNECT
CHECK
VALVE
PNEUMATIC
CONSOLE
433A
ARM
NO. 6
EDS FLIGHT
MONITOR
GAS HEAT
EXCHANGER
438A .
PNEUMATIC
CONSOLE
FILL MODULE
FROM
HELIUM
STORAGE
+4D15
LEGEND
CH.14384,2
Figure 5-38 5-37
Section V S-IVB Stage
and/or fire the APS engines. The control system for pitch and
yaw during the burn mode is essentially a linear, continuous type
whereas the APS during burn and coast is a nonlinear 'bang-bang'
system.
HYDRAULIC SYSTEM.
An independent closed-loop hydraulic system (figure 5-49) on the
S-IVB stage gimbals the J-2 engine during firing and non-firing
operations. A main hydraulic pump driven by the lox turbopumpturbine, a compressed air tank, an auxiliary motor-driven hydraulic
pump, an accumulator-reservoir, two servoactuator assemblies, and
interconnecting tube and hose assemblies comprise the hydraulic
system. The servoactuators, connected between the J-2 engine and
the thrust structure (one in the pitch plane and one in the yaw
plane), provide the forces necessary to gimbal the engine as com-
manded by the flight control computer. The actuators can operate
individually or together, either in phase or out of phase, to ac-
complish gimballing. The double-acting actuators are capable of
delivering 42,000 lbf at a pressure of 3650 psi in either the extendor retract direction.
Design Philosophy.
The philosophy of the design was to keep the system simple without
compromising reliability. The system is very similar to the hydraulic
system that was flown on the S-IV stages of the Block II Saturn
vehicles. The system features modular design to minimize the
number of external lines and reduce leakage paths. The auxiliary
motor-driven hydraulic pump, which is plumbed into the system,
provides an onboard checkout capability and provides a degree
of redundancy by operating in parallel with the engine-driven pump
during J-2 engine operation. A mechanical feedback loop and adynamic pressure feedback loop between the actuator and ser-
vovalve eliminate the necessity for an electrical feedback system.
Thermal conditioning features, such as a thermal isolator mount
for the engine-driven pump and circulation of hydraulic fluid
through the auxiliary pump motor jacket to distribute heat to the
hydraulic system components, prevent the system from freezing
during a ground hold.
Main Hydraulic Pump, Engine-Driven.
In the single stage, yoke type, variable displacement, pressurecompensated, axial piston, S-IVB main pump (figure 5-50), pump-
ing is accomplished by the reciprocation of nine pistons mounted
parallel to the cylinder block axis within the cylinder block. The
piston rods attach to the input drive shaft with ball sockets. The
cylinder block rotates synchronously with the input shaft, both
joined with a double universal joint. The rotating cylinder block
mounts in a yoke, which pivots to control the angular relations
between the cylinder block and drive shaft axes. When the axis
of the cylinder block aligns with the axis of the drive shaft (zero
angle), the pumping elements are in a zero-displacement or no-flow
position. As the yoke moves to increase the angle between the
shaft and the cylinder block, the pistons reciprocate over an increas-
ing stroke length, thereby increasing the displacement and the
output flow of the pump. A fixed valve plate, attached to the shaft
on which the cylinder block rotates, contains two ports, an inlet
port connecting the cylinder'suction with pump inlet and an outlet
port connecting cylinder discharge with pump outlet. As shown
in the main pump elements schematic, pistons at position A are
completely extended into the cylinder bores and pistons at position
B are completely withdrawn from the cylinder bores. As the cylinder
block rotates against the valve plate, each piston begins its suction
stroke at A, drawing fluid into the cylinder bore through the valve
plate inlet port. As the pistons pass position B, the discharge or
pressure stroke begins, forcing fluid from the cylinder bores through
S-IVB LOX TANK PRESSURIZATIONAND VENT SYSTEMDIAGRAM
CHECK__[
VALVE
CHECK j_
VALVE
PREPRE$SURIZATION
_LOOP
N\._ COLD
HELIUM Q-D
COLD HELIUMDUMP MODULE
NOTES:
[_ HELIUM PURGE, SEE S-IVBPNEUMATIC CONTROL SYSTEM
DIAGRAM
> 470 PSIG He SEE S-IVBt
PNEUMATIC CONTROL SYSTEM
DIAGRAM
LOX _,_---
J MODULE : (2 PtACES)
IREGULATOR-_ _3]!
J i
PL N M BACK UP
SHUTOFF !
Figure 5-3g
5-38
z
-_ ACTU AEIO N
I CONTROL
MODULE
(2 PLACES)
t
the valve plate outlet port; this continues until the piston reaches
position A where the cycle begins all over again.
Operation of the yoke type, variable displacement mechanism is
shown in the compensator schematic (figure 5-50). The pump yoke
is mounted on bearings, permitting the yoke to pivot and vary
the angle between the cylinder block axis and input shaft from
0 deg, zero displacement or no-flow to 30 deg, maximum dis-
placement or full-flow. The yoke is positioned by an actuating
cylinder mechanically coupled to an arm on the yoke. The yoke
spring retains the actuating cylinder in the maximum displacement
position under no-load conditions. During operations, this spring
force is opposed by control pressure fluid delivered by the compen-sator valve.
The compensator valve senses discharge pressure. When pump
outlet pressure exceeds 3650 psig, the hydraulic force on the com-
pensator valve spool overcomes the preset compensator valve springforce and displaces the spool downward, delivering high pressure
discharge fluid to the yoke actuating cylinder. This flow is propor-
tional to the compensator valve opening and to the excess discharge
pressure above 3650 psig. The yoke actuating cylinder reduces the
yoke angle until the flow is sufficient to maintain 3650 psig. The
compensator valve spool then centers to lock the yoke in the new
position.
A thermal isolator attaches the main pump to an accessory pad
on the lox turbine gas collector dome. The isolator protects the
pump, pump oil seal, and hydraulic fluid from temperature ex-
tremes effected by gases in the turbine that change suddenly from
-300 to +900 ° F during J-2 engine start. A dual element hot-gasshaft seal prevents impingement of hot gases on the hydraulic seal.
Any leakage past the seals is vented overboard. A crown-spline
quill shaft extending from the turbine shaft drives the main pump.
Compressed Gas Tank.
The compressed gas tank stores dry air to maintain an atmospheric
condition in the auxiliary pump motor case and to ensure proper
heat transfer from the motor to the hydraulic fluid circulating
through the motor case. The pressurized environment within the
motor also prevents excessive or rapid brush wear. A quick-discon-
nect fitting on the tank accommodates filling with air and a dial-in-
dicator type pressure gage provides local monitoring of tank pres-
sure. The tank is constructed of forged titanium and has a proof
pressure of 1200 "+ 25 psig.
Auxiliary Hydraulic Pump, Motor-Driven.
The single stage, electrically driven, variable delivery, fixed angle,constant-displacement pump (figure 5-51) s.upplies operating pres-
sure, hydraulic fluid, for preflight engine gimballing checkout, null
positioning during the boost phase, and emergency back-up during
S-IVB powered flight. As shown in the auxiliary pump schematic,view one, there is the piston and driver assembly, cylinder block,
and the valve plate. The pistons at position A are completely
extended into the cylinder bores and pistons at position B are
completely withdrawn from the cylinder bores. The pump accom-plishes pumping with the nine pistons attached by rods to the drive
plate at a fixed angle to the cylinder block axis. in view two, the
piston, the driver assembly, and the cylinder block are shown
connected. They rotate synchronously causing the pistons to recip-
rocate within their cylinder bores. The valve plate does not rotate
with the cylinder block. The pump output, hydraulic fluid flow
to and from the cylinder bores, is controlled by the position of
the valve plate. In view two (the maximum delivery position) as
the assembly rotates, fluid enters the cylinder bores during rotation
from A to B (suction stroke) and discharges during rotation from
B to A (discharge stroke). The valve plate inlet port connects each
cylinder to the pump inlet during the entire suction stroke, and
Section V S-IVB Stage
t0X SYSTEMMEASURMENTS
NUMBER I
XC4-403 I
VXCS-405 I
C16-424 I
C40-406 I
*C 159-424 I
C 163-424 I
C207-425 I
C208-405 J
C210-405 I
C2030-404 I
C2031-404 I
*XD3-403 J
*VXD 16-425 I
VDS5-424 J
VXD105-403 J
*VXD 179-406 J
*VXD180-406I
D219-403 J
D243-404 I
D244-404 J
XD248-42S I
D265-403 J
VXF4-424 I
*K2-424 I
K4-404 I
K16-404 I
K 18-404 I
K102-404 I
K 109-403 I
*KLI0-403 I
VK 156-404 J
K 198-424 I
K199-414 I
L4-406 I
L5.-406 I
NAME
i
TEMP, OXlD PUMP INLET
TEMP, COLD He SPHERE NO. 3 GAS
TEMP, OXlD TANK He INLET
TEMP, OXIDTANK POS I
TEMP, LOX CIRC RET LINE TK INLET
TEMP, OXlD CIRC PUMP OUTLET
TEMP, COLD He SPHERE NO. 5 GAS
TEMP, COLD He SPHERE NO. I GAS
TEMP, COLD He SPHERE NO. 4 GAS
TEMP, LOX NPV NOZZLE NO. 1
TEMP, LOX NPV NOZZLE NO. 2
PRESS, OXlD PUMP INLET
PRESS, COLD He SPHERE
PRESS, OXlD TANK INLET
PRESS, LOX TANK PRESS MOD He GAS
PRESS, OXIDTK ULLAGE EDS 1
PRESS, OXID TK ULLAGE EDS 2
PRESS, LOX CHILLDOWN PUMP DIFF
PRES_, LOX NPV NOZZLE NO. 1
PRESS, LOX NPV NOZZLE NO. 2
PRESS, COLD He SPHERES
PRESS, LOX PUMP INLET, CL COUPLED
FLOWRATE, OXID CIRCULATION PUMP
EVENT, OXlD TANK VENT VLV CLOSED
EVENT, OXID FILL VLV CLOSED
EVENT, OXfD TANK VENrr VLV 1 OPEN
EVENT, OXlD FILL VLV OPEN
EVENT, LOX PREPRESS FLIGHT SW ENABLED
EVENT_ OXID PREVALVE OPEN
EVENT, OXID PREVALVE CLOSED
EVENT, LOX TANK REO BACKUP PRESS ENABLED
EVENT, LOX NPV VALVE OPENED
EVENT, LOX NPV VALVE CLOSED
LEVEL, LOXTANK POS t
LEVEL, LOX TANK POS 2
RANGE
160 TO 170°R
25 TO 80°R
175 TO 560°R
160 TO 173°R
160 TO 320°R
163 TO 200°R
25 TO 80OR
25 TO 80°R
25 TO 560°R
100 TO 300°R
100 TO 300°R
0 TO 60 psia
0 TO 3500 psia
0 TO 300 psia
0 TO 500 psla
0 TO 50 I_ia
0 TO 50 ps;a
-30 TO 30 i_d
0 TO 50 psla
0 TO 50 psia
0 TO 3500 pda
0 TO 60 i_;a
0 TO 50 gpm
0TO5V
0TO5V
MEASUREMENT NUMBER PRERXES INDICATE THE FOLLOWING:
*PLIGHT CONTROL; V-ERE DISPLAY; X-AUXILIARY DISPLAY CH.20126
Figure 5-40
the valve plate outlet port connects each cylinder with the pump
outlet during the entire discharge stroke, thus the pump deliversfluid.
During operation, control pressure fluid metered by a compensator
valve opposes the compensator spring (as shown in the compensatorvalve schematic in figure 5-51). The compensator valve is sensitive
to discharge pressure and varies inversely with the flow. When
the pump discharge pressure exceeds 3650 psig, the compensatorvalve spool displaces, delivering high pressure fluid to the valve
plate vane. The hydraulic force acting on the valve plate vane
5-39
Section V S-IVB STage
PROPELLANTUTILIZATION SYSTEM
POWER ON
STATIC INVERTER
CONVERTER
I_LECTRONIC
PU ELECTRONICS
+4D 119
ENGINE
PNEUMATICS
NOTES:
_ SEE MEASUREh_NTS TABLE.
_ SWtTCH SELECTOR. SEEELECTRICAL SYSTEM.
•--------._ VXG 17
5-40
Figure 5-41
--IMIXTURE PATIO iCONTROL VALVE m
LOX TURBOUP MP __RJJ 2 ENGINE EL
TURBOPUMP
CH-201_
overcomes the compensator spring force and causes the valve plateto rotate until pump output is just sufficient to maintain 3650 psig.
In view three (the reduced delivery position) although the same
volume is displaced by the pistons, the valve plate ports a portion
of the discharge stroke (B to D) to the pump inlet and a portion
of the suction stroke (A to C) to the pump discharge. If the pump
outlet was completely restricted, the control fluid would rotate the
valve plate until (A-B) is perpendicular to (C-D).
In view four, the zero delivery position, each cylinder is open to
pump inlet and outlet during its entire suction and discharge stroke,
thus pump delivery is zero.
A regulator in the auxiliary pump manifold regulates the 475 psigair from the compressed air tank to 15 -- 5 psi for motor cavity
pressurization. The air serves as a heat transfer medium for thermal
conditioning of the hydraulic fluid circulating through the motor
jacket, and provides proper atmospheric conditions for the motor
operation.
The design safety factors for the auxiliary pump motor ensure
capable performance after operating at 125-percent speed at maxi-mum rated discharge pressure and capable performance with the
motor temperature of 150 ° F at start of cycle and 275 ° F at theend of a 5-min cycle.
Accumulator-Reservoir.
The hydraulic system accumulator-reservoir assembly (figure 5-52)
stores low pressure hydraulic fluid and provides pump ripple sup-
pression, pressure surge damping, and a source of high pressurefluid to supplement pump outputs during S-IB/S-IVB separation
transients. The accumulator-reservoir assembly consists of a mani-
fold assembly, a floating piston, a gas-loaded accumulator, a boot-
strap piston, reservoir, and reservoir piston. The manifold assemblycontains relief valves, bleed valves, quick-disconnects, a system
filter, and instrumentation. The accumulator is precharged with
GN 2 at 2350 ± 50 psig and 70 ° F. A dial indicator pressure gage
Section V S-IVB Stage
NAME _NGE
*VXG17-401
VK219-404
VXM1-411
VM4-411
VXM12-411
VM23-411
*VN1-411
*VXN2-411
*VN3-411
*VXN4-411
0 TO 65°POSITION, MRC VALVE
MRC VALVE OPEN ON
VOLT, STATIC INVERTER-CONVERTER 90 TO 135 V
VOLT, STATIC INVERTER-CONVERTER, 5VDC 4.5TO 5.5V
FREQUENCY, STATIC INVERTER-CONVERTER 390 TO 410 Hz
VOLT, STATIC INVERTER-CONVERTER, 21 VDC 20 TO 23.5 V
PU SYS LH 2 COARSE MASS, VOLT 0TO 5V
PU SYS LH 2 RNE MASS, VOLT 0TO 5V
PU SYS LOX COARSE MASS, VOLT 0 TO 5 V
FU SYS LOX RNE MASS, VOLT 0TO 5V
MEASUREMENT NUMBER PRERXES INDICATE THE FOLLOWING:
*FLIGHT CONTROL; V-ESE DISPLAY; X-AUXILIARY DISPLAY
CH.20121
Figure 5-42
on the accumulator provides local monitoring of the precharge
pressure. During pump operations fluid at 3650 "+ 50 psig entersthe accumulator. Accumulator filling continues until the floating
piston compresses the gas charge equal to the pressure of the fluid.
The high pressure fluid acts against the reservoir piston to maintain
low pressure in the reservoir. During non-operating periods of the
pumps, the precharge pressure forces the bootstrap piston against
the reservoir piston to keep the reservoir pressurized to 170 psi.
The low pressure fluid in the reservoir prevents cavitation in the
pumps during start and run operations.
A high pressure, full flow, cartridge type filter element filters the
IS-IVB ENGINE CUTOFF-- -- --d_Jk
LH 2 TANK LATCHING
RELIEF VALVE OPEN
LOX TANK NPV
VALVE OPEN
PU INVERTER & DC
POWER OFF
CSM/S-IVBSEPARATION
PASSIVATION
ENABLE
COLD HELIUM
DUMP
ENGINE CONTROL
HELIUM DUMP
STAGE CONTROL
HELIUM DUMP
-ll
---m
.... m
..... , ......... = ........
TIME FROM 5-1VB ENGINE CUTOFF (HR)
TNOTE: FOR EXACT TIME SEE
FIGURE 5-44.
i_i--/
|
m
CH-20122
Figure 5-43
5-41
Section V S-IVB Stage
COMMAND
S-IVB ENGINE CUTOFF NO, 1 ON
(START OF TIME BASE 4)
S-IV8 ENGINE CUTOFF NO. 2 ON
LOX TANK NPV VALVE OPEN ON
LH 2 TANK LATCHING RELIEF VALVE OPEN ON
PREVALVES CLOSE ON
CHILLDOWN SHUTOFF VALVES CLOSE
LOX TANK PRESSURIZATION SHUTOFF VALVES CLOSE
LOX TANK FLIGHT PRESSURIZATION SYSTEM OFF
PROPELLANT DEPLETION CUTOFF DISARM
MIXTURE RATIO CONTROL VALVE CLOSE
MIXTURE RATIO CONTROL VALVE BACKUP CLOSE
FLIGHT CONTROL COMPUTER S-IVB BURN /VK3DE OFF "A"
FLIGHT CONTROL COMPUTER S-IVB BURN MODE OFF "B"
AUX. HYDRAULIC PUMP FLIGHT MODE OFF
S/C CONTROL OF SATURN ENABLE
S-IVB ENGINE EDS CUTOFF DISABLE
LOX TANK NPV VALVE OPEN OFF
LOX TANK VENT AND NPV VALVES BOOST CLOSE ON
LOX TANK VENT AND NPV VALVES BOOST CLOSE OFF
P.U. INVERTER AND DC POWER OFF
LH 2 TANK LATCHING RELIEF VALVE OPEN OFF
LH2 TANK VENT AND L/R VALVES BOOST CLOSE ON
LH 2 TANK VENT AND L,/R VALVES BOOST CLOSE OFF
LH 2 TANK LATCHING RELIEF VALVE OPEN ON
LOX TANK NPV VALVE OPEN ON
LH 2 TANK LATCHING RELIEFVALVE LATCH ON
LOX TANK NPV VALVE LATCH OPEN ON
LH 2 TANK LATCHING RELIEF VALVE OPEN OFF
LOX TANK NPV VALVE OPEN OFF
LH 2 TANK LATCHING RELIEF VALVE LATCH OFF
LOX TANK NPV VALVE LATCH OPEN OFF
PASSIVATIO N ENABLE
LOX TANK PRESSURIZATION SHUTOFF VALVES OPEN
ENGINE He CONTROL VALVE OPEN ON
ENGINE He CONTROL VALVE OPEN OFF
PASSIVATIO N DISABLE
LOX TANK PRESSURIZATION SHUTOFF VALVES CLOSE
ENGINE PUMP PURGE CONTROL VALVE ENABLE ON
LOX TANK PRESSURIZATION SHUTOFF VALVES OPEN
ENGINE PUMP PURGE CONTROL VALVE ENABLE OFF
LOX TANK PRESSURIZATION SHUTOFF VALVES CLOSE
Figure 5-44
5-42
CHANh
12
48
42
99
82
91
79
104
98
6
35
iu/12
IU/5
29
Iu/18
IU/3
37
95
96
8
lOO
77
78
99
42
52
lOl
lOO
37
19
lO2
85
80
lO9
11o
86
79
24
80
25
79
0:00:0.1
0:00:.0.2
0:00:0.3
0:00:0.4
0:00:0.5
0:.00:.0.6
0:00:.0.8
0:00:1.0
0:00.1.8
0:00.2.2
0:00:2.4
0:00:.3.5
0:00:.3.7
0:00:3.9
0:00:5.0
0:00: TO.O
0:03: 20, 0
0:03:23.0
0:03:25.0
0:04:.00,0
0:05:20.0
0: 05:23.0
0:05:25.0
O: 15:(]0.0
0:15:0.2
0:15:2,0
0:.15:2,2
0:15:3.0
0:15:3.2
O:15:4.0
0:15:4,2
I: 23:10.2 :
1:23:20.0 J
1:23:,20.2 1
1:40:00.0 I
1:40:.0.1 I
!: I_.00,0 J
_.0_.0 I
[:20:00.0 I
k30:00.0 J
k,36:,40.0 I
CH-2017ql
fluid entering the accumulator reservoir from the pumps. The
15-micron filter has an operating pressure of 3700 psig and a
differential pressure of 15 psi maximum at 12 gpm and 100 --3 ° F.
A high pressure relief valve prevents fluid over-pressurization in
the accumulator by relieving excessively pressurized fluid into the
reservoir. The valve cracks at 3760 psi, is fully open at 4250 psi,
and reseats at 3620 psi. The valve has a proof pressure of 5600psi.
Two balanced, low pressure relief valves connected in series protect
the low pressure reservoir and fluid return side of the system. Both
valves have a cracking pressure of 275 psig at -70 to + 110 ° F,
a relief pressure of 290 psig, and a reseat pressure of 225 psigminimum at 120 ° F. Instrumentation transducers are covered in
Hydraulic System Measurements.
Hydraulic Servoactuators.
The S-IVB actuator (figure 5-53) consists.of a tailstock, a body-
cylinder member, a piston, and a front bearing member. The
actuator attaches to the thrust structure with a spherical bearinghoused in the tailstock.
The two self-contained, equal area, double acting actuators convert
hydraulic power into a mechanical output that controls engine
movement and vehicle attitude in two perpendicular planes (pitchand yaw). The actuator body flange mounts to the tailstock. The
forged aluminum body houses the pressure and return ports, theelectrical connectors, the prefiltration valve, the servofilter, the
cylinder bypass valve, bleed valve, and the servovalve piston. Ser-vovalve and telemetry potentiometers are mounted to the actuator
body.
The servovalve is a four-way, two-stage, flow control valve with
a dynamic pressure feedback loop. The valve consists of a polarized,
dry coil, electrical torque motor and two stages for hydraulic ampli-
fication. The polarizing magnetic flux is generated by two perma-
nent magnets mounted between the upper and lower pole pieces.
The motor armature extends into the air gaps of the magneticflux, supported in this position by the flexible tube member. Two
motor coils surround the armature, one located on each side of
the flexure tube member. The flexure tube member also acts as
a seal between the electromagnetic and hydraulic sections of theservovalve.
The flapper of the first hydraulic amplifier stage is rigidly attached
to the midpoint of the armature. The flapper extends through theflexure tube member and passes between two nozzles in the control
pressure loop and two nozzles in the dynamic feedback loop, creat-
ing two variable orifices between the nozzle tips and the flapper.
Filtered fluid from the pressure source is supplied to the variable
orifices in the control pressure loop through two fixed upstreamorifices. Pressure from a point between the fixed and variable
orifices is transmitted to the ends of the spool in the second hydrau-lic amplifier stage.
When a signal is applied to the motor windings, the flapper moves
toward one of the nozzles (which nozzle depends upon the polarityof the signal). The flapper movement toward one nozzle tends to
restrict fluid flow through that nozzle which results in a pressure
build-up that acts against the spool. The opposite nozzle has less
restriction and a greater fluid flow results in a lower pressure onthe opposite end of the spool. This pressure differential causes the
spool to reposition. In doing so, one side of the spool aligns high
pressure supply with one of the actuator fluid loops while the other
side of the spool ports fluid from the actuator to the low pressurereturn passages.
The second stage spool is a conventional four-way, sliding design
Section V S-IVB Stage
HELIUM FROM
STORAGE FACILITY
/_PNEUMATIC
-- /CONSOLE-- /432A
oooo
z!
e..io
UNION _m
VALVE
_ P,ACES>_
91
81
Figure 5-45 (Sheet 1 of 2) 5-43
Section V S-IVB Stage
S-IVB PNEUMATICCONTROLSYSTEM
!
I!
I!
I-|
LH 2 VENT
LH 2 VENT DIRECTIONALCONTROL
FLIGHT
OPEN\LATCHINGLvE i
BOOST CLOSE
...=..=LATCH
It
TANK VENT &
RELIEF VALVE_smm...n...|m...m.,||.m.m.mm|.|.ll
_..-...._ . ....I
I "
SHUTOFF VALVE
J _...............
r roB-- CHECK i| VALVE (18 PLACES)
(8 PLACES)
LEGEND
32(30 OR 1000 PSIG HELIU/vL
475 PSIG HELIUM DIST_B_ION,_ ai_
VALVE ACTUATION HELIUM,,,,,m,,,,,,.Nm,,.m,
PURGE HELlUMm i .lB. _ ,i i m _ i roll=
ELECTRICAL SIGNAL
!
\• FROM
SHEET 1
5-44
i
Figure 5-45 (Sheet 2 of 2)
NOTES
[_::> SEE MEASUREMENTS TABLE
[_> SWITCH SELECTOR.
SEE ELECTRICAL SYSTEM
C-H 14308-1
SectionV S-IVB Stage
S-IVB PNEUMATIC CONTROL SYSTEM DIAGRAM
/--- LOX SENSE
1.1,=/ LINE PURGE
H: :___,/--LO×NPV_CTPORGEI__ _'-- LOX VENT
& RELIEF HELIUM
VALVE PURGE STORAGE
SPHERE _
I [_] _[_ _ LH2 CHILLDOwN VALVE PURGE _
PURG
PURGE /(TYP)
_-.LH 2NPV
DUCT
PURGE
NOTES:
_SEE S-IV8 FUEL SYSTEM DIAGRAM
> SEE S-IVB FUEL TANK PRESSURIZATIONAND VENT SYSTEM DIAGRAM
SEE S-IVB LOX SYSTEM DIAGRAM
SEE S-IV8 LOX TANK PRESSURIZATIONAND VENT SYSTEM DIAGRAM
T.F I F.NUM
PNEUMATIC REGULATOR
CONTROL ASSEMBLY
_;_:U__v,,,, _ _iLH 2 LH2 FILL FILL ANDDIRECTIONAL AND DRAIN DRAIN
CONTROL VALVE VALVE VALVE
LH2 VENT & RELIEFVALVE (OPEN)
VENT & RELIEF &
LATCHING VALVE BOOST CLOSE
_ ,
LH 2 LATCHING LOX VENT & RELIEF LOX NPVRELIEF VALVE VALVE (OPEN} VENT VALUE
(OPEN & LATCH) & RELIEF & NPV BOOST CLOSE (OPEN& LATCH)
LOX CHI LLDOWN
MOTOR CO NTAI NER
PURGE
....PUMP PURGE CHILLDOWN (CLOSE) J-2 ENGINE START
CONTROL VALVES (CLOSE) TANK VENT & RELIEF
MODULE VALVE CONTROL
-1
_J
CH.14716-1
Figure 5-46 5-45
Section V S-IVB Stage
PURGED COMPONENT FLOW RATE
LH 2 VENT DUCT
LOX NPV DUCT
LOX FILL AND DRAIN VALVE
LH 2 FILL AND DRAIN VALVE
FUEL CHILLDOWN SHUTOFF VALVE
LOX VENT AND RELIEF VALVE
LOX CHILLDOWN PUMP
LOX ULLAGE SENSI NG LI NE
ENGINE TURSOPUMPS AND GAS GENERATOR
Figure 5-47
1728 SCIM
1728 SCIM
15 SCIM
15 SCIM
24,200 SCIM
112,300 SCIM
600 SCIM
1728 SCIM
16,400 SCIM
Ct1.14511.1
in which output flow from the valve at a fixed pressure drop is
proportional to spool displacement from the null position. A can-
tilever feedback spring is fixed to the armature and extends through
the flapper to engage a slot at the center of the spool. Displacement
of the spool deflects the feedback spring, creating a restoring torque
on the armature. The output flow from the valve is ported to either
side of actuator piston, causing the piston to extend or retract andmoving the required load.
A major feature of this servovalve is its dynamic pressure beedback
(DPF) loop. This loop dampens gain peaking under dynamic condi-tions, providing the necessary stability. Under static conditions the
feedback loop is ineffective in eliminating positional errors due
to static loads. The feedback loop consists of a spring-centered,
DPF piston connected to a second pair of nozzles located on either
side of the flapper. Each chamber at the end of the DPF pistonis open to an actuator chamber; therefore, differential pressures
across the actuator, act on the ends of the DPF piston, causingthe piston to seek a new equilibrium position and to displace a
quantity of hydraulic fluid in the process. The fluid flows through
one of the DPF nozzles and impinges against the flapper, providinga feedback force proportional to actuator forces. (This feedback
condition occurs only under dynamic loads; under static loads the
DPF piston ahcieves the new equilibrium position as forces on
the centering spring equalize, ceasing feedback fluid flow.)
A mechanical feedback device, which consists of a lanyard con-
nected between the torque motor armature and roller-ramp assem-
bly helps maintain the servoactuator assembly in a null position
by keeping the flapper centered in the servovalve. A mechanical
adjustment on the armature permits alignment of the servoactuator
null position.
A cartridge type, full flow, vented seal, stainless steel mesh filter
unit filters all operating fluid before it reaches the servovalve. Fluid
passages are so designed that direct impingement of fluid uponthe filter element is avoided. The element is rated at 5 microns
nominal, 15 microns absolute, both at 1.0 gpm flow.
A manually operated, prefiltration valve bypasses all actuator com-
ponents by connecting the supply pressure port with the return
pressure port. The prefiltration valve permits flushing of the hy-
draulic system without exposing sensitive components of the actua-
tor to the contaminated fluid. The servovalve (bypass valve) is
a manually operated, normally closed valve that interconnects the
actuator ports, allowing manual movement of the actuator. A
midstroke locking device provides a means for actuator installation
without accidental actuator piston movement and prevents random
engine movement during servicing, shipping, and storing.
There are several bleed valves in the system. Bleed valves are
connected to the cylinder ports and to the servovalve operating
pressure fluid inlet passage. The cylinder bleed valves allow gas
or hydraulic fluid to vent to atmosphere, also the servovalve bleed
valve allows samples of filtered operating fluid to be taken.
Hydraulic System Servicing.
The hydraulic system servicing consists of assuring the system is
charged with approximately 12 Ibm of MIL-H-5606 hydraulic fluid,
and the accumulator is charged with GN z to the equivalent of
2350 -+- 50 psig at 70 ° F. The total quantity of the accumulator
GN 2 will be 1.9 -+- 0.1 Ibm. The auxiliary pump air supply tank
HYDRAULICSYSTEMINSTL
COMPRESSED
GAS RESERVOIR
NUMBER NAME
*VXD14-403
*VD5D-403
VD 103-403
*VXD160-403
*XD247-403
*XD255-403
VK 105-404
PRESS, CONTROL He REG DISCHARGE
PRESS, ENGINE PUMP PURGE REG
PRESS, He PRESS TO LOX MOT CNTR
PRESS, AMBIENT He SPHERE
PRESS, CONT He REG DISCHARGE
PRESS, AMBIENT He SPHERE
EVENT, PUMP PURC.E REG BACKUP DEENERGIZED
MEASUREMENT NUMBER PREFIXES INDICATE THE FOLLOWING:
*PLIGHT CONTROL; V-ESE DISPLAy; X-AUXILIARY DISPLAY
Figure 5-48
5-46
RANGE I
0 TO 650 mi_ I
0 TO 150 p_ia I
0 TO 3500 pslaJ
0 TO 650 psla [
0 TO 3500 psla]
CH-20128
AUXILIARY
HYDRAULICPUMP
MAIN
HYDRAULIC
PUMP
Figure 5-49
C.H 14560
Section V S-IVB Stage
LOCATION J-2 ENGINE, OXIDIZER TURBINE EXHAUST DUCT
SERVICE HYDRAULIC FLUID MIL-H-5606
OPERATING PRESSURE
OUTLET PRESS.
INLET PRESS.
CASE PRESS.
INLET & CASE PRESS.
PROOF PRESSURE
OUTLET PRESS.
INLET PRESS.
CASE PRESS.
BURST PRESSURE
OUTLET PRESS.
INLET PRESS.
CASE PRESS.
OVERSPEED
TEMPERATURE
AMBIENT
FLUID
RATED FLOW
STARTING TORQUE
RESPONSE
3650 psig MAX.
190 i_;g MAX.
190 psig MAX.
320 pdg MAX. - SERVICING (NON-OPERATING)
5475 psig MIN.
480 pslg MIN.
480 pslg MIN.THE PUMPS SHALL MEET ALL REQUIREMENTS AFTER
SUBJECTION TO THE ABOVE PRESSURES FOR 3
MINUTES AT 275°F.
9t25 pdg MIN.
800 psig MIN,
800 pslg MIN.THE PUMPS SHALL MEET ALL REQUIREMENTS AFTER
SUBJECTION TO THE ABOVE PRESSURE FOR 3
MINUTES AT 275°F WITH NO RUPTURE OR STRUCTURE
FAILURE.
THE PUMPS SHALL MEET ALL REQUIREMENTS
WITHOUT DAMAGE AFTER OPERATING AT
10,000 rpm's FOR 30 MINUTES.
-35 TO +275°t:
-35 TO +275°/:
7 gpm AT THESE OPERATING CONDITIONS:
SPEED MAX. 7000 rpm
OUTLET PRESS. MIN. 3550 psig
INLET PRESS. MAX. 150 pdg
THE MAXIMUM ALLOWABLE TORQUE TO ACCELERATE
THE PUMP FROM STATIC CONDITION TO 9000 rpm
LINEARLY IN 1.5 =e¢. WITH STEADILY APPLIED 3550
psig OUTLET PRESSURE AND 150 pdg INLET PRESSURESHALL NOT EXCEED 220 ;n.-II_. AFTER 8 hrs. SOAK
pERIOD AT 0°F.
THE PUMP FLOW COMPENSATOR SHALL RESPOND
WITHIN 50 ms a) FROM 80% FULL FLOW TO 5% AND
b) FROM 5% FULL FLOW TO 80% FULL FLOW.
YOKE-TYPE VARIABLE DISPLACEMENT MECHANISM SCHEMATIC
YOKE
ACTUATING
CYLINDER _'-_ I
PUMP
INPUT
SHAFT
PUMP YOKE J
_- YOKE SPRING
0 _ ,
_/'- CASE_
I PRESSUREADJUSTMENT "_'_
' PUMP OUTLET
COMPENSATOR
VALVE SPOOL
' PUMP INLET
Figure 5-50
receives an initial charge equivalent to 475 psig at 70 ° F prior
to propellant loading operations.
Hydraulic System Operation.
The S-IVB stage hydraulic system performance is monitored at
the S-IVB hydraulics panel during countdown and launch opera-
tions. Switches and indicators on the hydraulics panel provide
control and monitoring of system pressures, temperature, and hy-
draulic oil level. During propellant loading, the auxiliary hydraulic
pump is cycled on from the hydraulic panel using the AUX HY-DRAULIC PUMP POWER switch when the MAiN PUMP INLET
OIL TEMP indicates - 10 ° F minimum and remains on until the
.FIXED
VALVE
plATE
12J
VARIABLE DISPLACEMENT
MECHANISM VARIES ANGLE
-- DIgCTHARGE OF DRIVE ASSEMBLY'_
f SUCTION PORT A --_,_
CYLINDER PISTON AND DRIVER
BLOCK ASSEMBLY
11 10_9 8 7
1. THERMOCONDITIONING FLUID TO AUXILIARY PUMP
2. VALVE, HYDRAULIC BLEEDER
3. VALVE, CHECK, HIGH PRESSURE
4. FLUID DISCHARGE
5. PUMP, HYDRAULIC, ENGINE DRIVEN6. FLUID RETURN
7. SWITCH, THERMAL, MAIN PUMP INLET8. TRANSDUCER, TEMPERATURE, MAIN PUMP INLET
9. VALVE, CHECK, HIGH PRESSURE
10. HYDRAULIC PUMP, PIPE PLUG ASSEMBLY
11. THERMAL ISOLATOR
12. DRIVE SHAFT C-H 1421_
temperature indicates approximately 80 ° F. The presence of the
auxiliary pump on command illuminates the AUX HYDRAULIC
PUMP POWER ON light on the panel. The auxiliary pump is
programmed to operate continuously through the S-IVB switch
selector in the flight mode from approximately T-11 min prior to
launch, throughout the boost phase and S-IVB engine operation.
In the event the reservoir oil level reaches approximately 10%,
the RESERVOIR OIL LEVEL LOW talkback light will illuminate.
In the event the inlet oil temperature should reach -15 ° F, the
MAIN PUMP INLET TEMP LOW talkback light will illuminate.
During normal system operation, low pressure fluid from the ac-
cumulator-reservoir flows through the main engine-driven pump
5-47
Section V S-IVB Stage
AUXILIARY MOTOR-DRIVENHYDRAULICPUMF
LOCATION J-2 ENGINE THRUST STRUCTURE
SERVICE ELECTRICAL & HYDRAULIC FLUIDMIL-H-5606
OPERATING PRESSURE
OUTLET PRESS.
INLET PRESS.
CASE PRESS.
PROOF PRESSURE
OUTLET PRESS.
INLET PRESS.
CASE PRESS.
BURST PRESSURE
OUTLET PRESS.
INLET PRESS.
CASE PRESS.
TEMPERATURE
AMBIENT
FLUID
ELECTRICAL MOTOR
TYPE
VOLTAGE
STARTING CURRENT
RUNNING CURRENT
RATED FLOW
RESPONSE
3650 psig MAX.
45 pslg MIN.
45 pslgAT 275OF
5475 pslg MIN.
480 psig MIN.
480 pslg MIN.at 275°F
9125 pslg MIN.
800 pslg MIN.
800 psl9 MIN.
-30 TO +275OF
0 TO +275°F
CONTINUOUS DUTY SHUTFIELD
50 TO 60 vdc
250 A
75 A
1.5 gpm
100 ms UNDER NORMAL
OPERATING CONDITIONS.
1. FILTER
2. VALVE, RELIEF
3. HYDRAULIC PUMP PIPE PLUG ASSEMBLY
4. DETECTOR, CHIP
5. VALVE, HYDRAULIC BLEEDER
6. VALVE, RELIEF
7. VALVE, CHECK
8. RETURN FROM ACCUMULATOR
9. DISCHARGE TO ACCUMULATOR
10. RETURN TO MAIN PUMP
I1. DISCHARGE FROM MAIN PUMP
12. VALVE, CHECK
13. COUPLING HALF, QUICK DISCONNECT
14. HYDRAULIC PUMP PIPE PLUG ASSEMBLY
15. REGULATOR, AIR PRESSURE
16. VALVE, CHECK17. AIR FROM DRY AIR BOTTLE
18. THERMOCONDITIONING FLUIDFROM MAIN PUMP
19. VALVE, RELIEF
OiT/___O'o'C#ARGESUCTION PORT A
, _VALVE CYLINDER PISTON AND DRIVER
PLATE BLOCK ASSEMBLY
VIEW I. COMPONENT SCHEMATIC
19 ' LEGEND I
r--1 HIGH-PRESSURE FLUID
475 PSIG AIR
B I5 PSIG AIR
i _ LOW PRESSURE FLUID
_ CONTROL PRESSURE FLUID
I
SPOOL'_ /'-- COMPENSATOR
PUMPOOTLJI PR'NGPRESSUR,---__iiiiiiii_
•_f t _ TOCASE
DRAIN
TO VALVE PLATE VANE
C A_
VIEW 3. REDUCED DELIVERY POSITION
5-48
D
VIEW 2. MAXIMUM DELIVER'_ POSITION
Figure 5-51
VIEW 4. ZERO DELIVERY POSITION
C.H 14275
Section V S-IVB Stage
14 13
15 10
8 7
19 20
Figure 5-52
case and enters the auxiliary pump motor jacket for thermal condi-
tioning. The fluid then flows through a filter and enters the auxiliary
hydraulic pump. A relief valve will permit the fluid to bypass the
filter if the differential pressure exceeds 90 psid. A check valve
in the low pressure auxiliary pump inlet will permit return flow
of main pump case fluid if the auxiliary pump should fail. The
pump increases the fluid pressure to 3,650 psig and delivers the
fluid through a pump discharge check valve to the accumulatorreservoir.
The hydraulic fluid flows through the main system filter element
into the accumulator chamber where the hydraulic volume increases
until a fluid-to-gas pressure balance equivalent to the discharge
pressure is obtained. The fluid exits the accumulator and flows
to the pitch and yaw servoactuators. The fluid flows through anactuator filter into the servovalve. Inside the servovalve, the fluid
is filtered again before it enters the critical control passages. With
the servovalve commanded to the null position, the fluid flows
through the variable orifices into the low pressure return passages,
which port the fluid back to the accumulator-reservoir. Flowrate
in the null mode of operation is 0.4-0.8 gpm, which is the totalof both servovalve leakage rates.
With the system pressurized to full operating pressure by the auxil-
iary pump at J-2 engine start, the main pump compensator moves
the main pump yoke toward the no-flow position, thereby reducing
LOCATION J-2 ENGINE THRUST STRUCTURE
SERVICE HYDRUALIC FLUID MIL-H-5606
TEMPERATURE -12 TO +275°F
PRESSURE
ACCUMULATOR 2350 ± 50 pda GN 2RESERVOIR
LOW 170 F4ig
HIGH 3650 psig
1. TRANSDUCER, PRESSURE, GN 2 ACCUMULATOR
2. TRANSDUCER TEMPERATURE, GN 2 ACCUMULATOR3. DISCHARGE TO YAW ACTUATOR
4. DISCHARGE TO PITCH ACTUATOR
5. VALVE, HYDRAULIC BLEEDER
6. RETURN FROM BOTH ACTUATORS
7. VALVE, HYDRAULIC BLEEDER
8. TRANSDUCER, PRESSURE, RESERVOIR OIL
9. POTENTIOMETER, PISTON POSITION
10. TRANSDUCER, TEMPERATURE, RESERVOIR OIL
11. TRANSDUCER, PRESSURE, HYDRAULIC SYSTEM
12. VALVE, HYDRAULIC BLEEDER
13. VALVE, LOW PRESSURE, BALANCED
|4. FITTING, HOSE ADAPTER
15. LOW-PRESSURE FLUID RETURN TO PUMP
16. HIGH-PRESSURE FLUID FROM PUMP
17. COUPLING HALF, QUICK DISCONNECT
18. VALVE, RELIEF, LOW PRESSURE BALANCED
19. VALVE, RELIEF, HIGH PRESSURE
20. COUPLING, HALF, QUICK DISCONNECT
21. FILTER ELEMENT, MAIN SYSTEM
22. VALVE, RELIEF, VENT, LOW PRESSURE
23. GAGE PRESSURE, DIAL INDICATOR
24. VALVE, AIR, HIGH PRESSURE, CHARGING
25. VALVE, RELIEF, VENT, LOW PRESSURE
26. ACCUMULATOR
27. LOW-PRESSURE RESERVOIR
28. HIGH-PRESSURE RESERVOIR
29. FLOATING PISTON
30. RESERVOIR PISTON
31. BOOTSTRAP PISTON
C-H 14276
starting torque requirements upon the turbine-driven quill shaft.
As the pump accelerates, the compensator increases the pump-yoke
angle which controls the output. When the pump reaches full speed,
its discharge flow adds to the auxiliary pump flow and the hydraulic
system is in a state of readiness. Check valves in the discharge
lines of the pumps prevent pressure reversal whenever the system
is pressurized with the pumps inactive, and prevent interaction
between the pumps during their operation.
Hydraulic System Fluid Requirements Summary.
Hydraulic Fluid. The hydraulic fluid shall meet the requirements
of MIL-H-5606A. The properties of the petroleum base fluid are:
fluid temperature range of-65 ° F to +275 ° F for closed system;
pour point of less than -90 ° F; flash point of +230 ° F.
The solid contaminant particles content shall not exceed the allow-
able number of particles per 100 ml sample of fluid: 1340 particles
0-25 microns in size; 210 particles, 25-50 microns in size; 28
particles, 50-100 microns in size; 3 particles, including fibers, over100 microns in size.
Foaming characteristics of the hydraulic fluid when tested at 75 °
shall be: stable foaming tendency and complete foam collapse after
a 5-minute blowing period followed by a 10-minute settling period.
A ring of small bubbles around the edge of the graduate constitutes
5-49
SectionVS-IVBStage
complete collapse. A random sample of filled unit containers and
a sample of shipping containers fully prepared for delivery shall
be selected from each lot of the hydraulic fluid.
Air. Air used for filling the auxiliary hydraulic motor compressed
gas tank conforms to the following specification: In a 30-ft _ volume
of air the maximum allowable number of particles 30 to 100 microns
in size is 25 and no particles over 100 microns in size are allowed;
the total allowable hydrocarbon content is 25 ppm expressed by
equivalent weight of carbon; the maximum allowable moisture
content is 25 ppm by volume (equivalent to a dew point of -64 °F).
Gaseous Nitrogen. GN 2 used for the accumulator-reservoir pre-
charge conforms to MSFC-SPEC-234. The purity of GN 2 is 99.98percent (min) nitrogen by volume.
Oxygen content is 150 ppm (max) by volume, hydrocarbon content
(expressed as methane) is 15 ppm (max) by volume, and moisture
content is 11.5 ppm (max) by volume at standard conditions.
InterconnecUon Plumbing.
Ja-.th high and low pressure flexible hoses are constructed of poly-
tertrafiuoroethylene (teflon) lined hose reinforced with steel wire
braid. The flex hoses are normally straight (not precurved) and
use straight fittings.
The hydraulic tube assemblies are constructed of steel tubing cor-
rosion-resistant (CRES) 304, aerospace vehicle hydraulic system1/8 hard condition.
Critical Components.
All the hydraulic system components are flight critical. They are:
servoactuators; main hydraulic pump, engine-driven; auxiliary hy-
draulic pump, motor-driven; and accumulator-reservoir.
Hydraulic System Measurements.
Nine measurements (figure 5-54) are utilized to monitor systemreadiness condition and operation. These measurements are tele-
metered through the S-IVB stage telemetry system to ground re-
ceiving stations where they are monitored in real-time in addition
to being recorded. Seven meters on the S-IVB hydraulics panel
provide monitoring for each measurement. Pitch and yaw actua-
tor-position measurements are monitored on the S-IVB engine
deflection panel meters. Flight control measurements in figure 5-54are relayed to mission control center for real-time monitoring.
AUXILIARY PROPULSION SYSTEM.
The auxiliary propulsion system (APS) provides attitude control
for the S-IVB stage/IU and payload during both powered and
coast phases of flight. During the powered phase, the APS providesonly roll control; gimballing the J-2 engine provides yaw and pitch
control. During the coast phase, the APS provides pitch, yaw, and
roll control (figure 5-55). The APS configuration (figure 5-56)
consists of two aerodynamically shaped modules, each 80 in. high,
installed on the aft skirt 180 deg apart at positions I and III. Each
module contains three liquid-propellant rocket engines, a fuel and
oxidizer storage and supply system, a high pressure helium storage
and regulator system, and control components for preflight servicing
and infiight operation.
The APS uses nitrogen tetroxide (N 2 O 4) as oxidizer, and mono-
methylhydrazine (MMH) as ihe fuel. The fuel and oxidizer are
hypergolic; therefore, an ignition system is not required. Attitude
HYDRAULICACTUATORASSEMBLY
LOCATION
SERVICE
PISTON STROKE
PRESSURE
OPERATING
PROOF
BURST
ELECTRICAL
SERVOVALVE
MOTOR CURRENT
RESISTANCE
FEEDBACK POT
VOLTAGE
RESISTANCE
ATTACHED BETWEEN J-2
ENGINE & THRUST STRUCTURE
ELECTRICAL & HYDRAULIC FLUID
MIL- H-5606A
2.90 _*0.032 in. 2
3650 pslg
5475pslg_ , 3650psig9130 p$ig <_ , 5475 psig ._]
50 mA
I00_+ 5fl
60 Vde
2000 ± 20 fl
I_FLUID PRESSURE SECTION
2_FLUID RETURN SECTION
LEGEND
J_ HIGH PRESSURE SUPPLY
_1 LOW PRESSURE RETURN
DYNAMIC PRESSURE FEEDBACK LOOP
_ CONTROL PRESSURE LOOP
ACTUATION FLUID LOOP
5-50
I. ACTUATOR
2. VALVE, CYLINDER BYPASS
3. VALVE, HUDRAULIC BLEEDER
4. FILTER, SERVOVALVE
5. VALVE, PREFILTRATION
6. DISCHARGE FROM RESERVOIR
7. RETURN TO RESERVOIR
8 SERVOVALVE
9. POLARIZED TORQUE MOTOR
10. ARMATURE
11. MECHANICAL ALIGNMENT ADJUSTMENT
8 9 10 7 6
17
Figure 5-53
21
12. FLAPPER
13. CANTILEVER SPRING14. SPOOL
15. VARIABLE ORIFICES (TYP)
16. FIXED ORIFICES
17. MECHANICAL FEEDBACK LOOP
18. DYNAMIC PRESSURE FEEDBACK PISTON
19. ACTUATOR PISTON
"20. TAILSTOCK
21. BODy CYLINDER
22. FRONT BEARING MEMBER
4 3
22
CH-14302-t
corrections are made by firing the 150-1bf engines, individually
or in any combination, in short bursts of approximately 65 msec
minimum duration. Commands from the flight control computer
through the attitude control relay modules actuate fuel and oxidizer
solenoid-valve clusters that admit propellant to the engine combus-
tion chambers. Helium pressure exerted against stainless steel bel-
lows assemblies, which contain the fuel and oxidizer, forces propel-
lant into the engine. For a detail description of the engine actuation
control, see Section VI (Navigation, Guidance, and Control).
Attitude Control System Concept.
A detachable modular subsystem concept for the attitude control
system was selected for the S-IVB stage to maintain an independent
propulsion subsystem that facilitated development, qualification
checkout, manufacturing, and replacement if required. Also, the
modular concept is appropriate for isolating the thermal control
problems associated with space envrionments and the cryogenicspace vehicle systems. The extended orbital operation of the S-IVB
stage requires thermal protection of the hypergolic propellant sys-tems to prevent feeezing and overheating. Compact system modules,
with short propellant lines and buried engine installations greatly
simplify the temperature control problem by permitting the use
of a passive protection. Conventional insulation and control of the
surface emissivities are the only thermal protection required.
The attitude control system module was developed and qualified
as a system at the Sacramento test center. Firing tests verified
the interaction performance and operational characteristics of the
pressurization, positive expulsion propellant, and engine systems.
Also, verification of the checkout and operational procedures was
accomplished. The verification of thermal capabilities was accom-
plished in the 39-ft diameter space chamber at the McDonnell
Douglas space systems center. In addition, the modular concept
allowed a complete system vibration test to verify the installationsunder this environment.
The module was designed to operate with a 3000-psi helium supply
and a 196-psia propellant expulsion system. This allows the engine
to operate with a 100-psia chamber pressure at 150 lbf. The 3000-psi
helium pressurization system was chosen to be consistent with the
requirements for the main stage, which also utilizes helium at 3000
psi. The requirement for 196 psia was based on an optimization
weight study which provides the minimum weight tank system,
and still provided adequate pressure for operating the engines at
a 100-psia chamber pressure. A chamber pressure of 100 psia was
considered to be consistent with the requirements of a 150-1bf
engine for a satisfactory engine envelope and performance. The
thrust requirement was established based on the maximum thrust
required to control the S-IVB stage.
Hypergolic bi-propellants are utilized since they provide the advan-
tages of storability arrd spontaneous ignition; hypergolic ignition
allows a more simple engine design with a fast response rate. The
quantity of propellant required was established as 61 Ibm (nitrogen
tetroxide and monomethylhydrazine) at an oxidizer to fuel ratio
of 1.60. This quantity was based on a system study of the mission
requirements for attitude control and vehicle stabilization require-ments, with consideration of engine performance. The requirements
also include a need for high-response-rate engine valves to provide
a minimum impulse bit. It was determined that a total impulse
of 7.5 _ 0.75 lb-sec is satisfactory without significant degradation
of engine performance. This assumes that all propellant valves on
each engine will operate at a maximum pulse rate of 10 starts/sec.
Each engine has a maximum design life of I0,000 starts with a
maximum pulse rate of 10 starts/sec.
APS Major Subsystems.
Each of the two modules is comprised of three major subsystems:
Section V S-IVB Stage
I HYDRAULIC SYSTEM MEASURMENTS
NAME
*VXC50-401 TEMP, HYD PUMP INLET OIL
*VXC51-403 TEMP, RESERVOIR OIL
VC 138-403 TEMP, ACCUMULATOR GN 2
*VXD41-403 PRESS, HYDRAULIC SYSTEM
*VXD42-403 pRESS, RESERVOIR OIL
*VXD43-403 PRESS, GN 2 ACCUMULATOR
*VXG1-403 POSITION, ACTUATOR PISTON POT, PITCH
*VXG2-403 POSITION, ACTUATOR PISTON POT, YAW
* VXLT-403 LEVEL, RESERVOIR OIL
RANGE
400 TO 785°R
40O TO 785°R
400 TO 785°R
1500 T'O 450O PSIA
0 TO 400 PSIA
1500 TO 4000 PSIA
7.5TO -7.5 DEG
i -7.5TO 7.5 DEG
0 TO 100%
MEASUREMENT NUMBER PREFIXES INDICATE THE FOLLOWIIqG:
*FLIGHT CONTROL; V-ESE DISPLAY; X-AUXILIARY DISPLAY
CH-20124
Figure 5-54
helium pressurization, positive expulsion propellant feed, and atti-
tude control engines. The helium pressurization subsystem supplies
and regulates the flow of pressurized helium to the positive expul-
sion subsystem (propellant tankage). The positive expulsion subsys-
tem provides propellant (fuel and oxidizer), on demand, to the
engines under zero and random gravity environment. The attitude
control engines provide the necessary thrust for attitude controland maneuvering of the stage during various phases of flight.
Instrumentation on the module consists of the various pressure
and temperature readings for performance evaluation. A description
of each of these subsystems and their regulated components is
presented in the following paragrpahs.
Helium Pressurization Subsystem. The pressurization subsystem
consists of a fill disconnect, two check valves in series, a helium
storage tank, and a pressure regulator. Downstream from the regu-lator, the system splits into fuel and oxidizer pressurant branches.
Each branch includes a quad-che(:k value and a low pressure
module. During prelaunch operations, the 268 cu in helium storage
tank is pressurized through the fill disconnect and two check valves
to 3100 ± 100 psig. The helium tank incoprorates a temperature
probe for monitoring tank temperature. Tank pressure is measured
at the inlet to the helium pressure regulator.
Helium stored in the tank at 3100 -+- 100 psig is supplied to the
helium pressure regulator assembly. The helium gas entering the
APS ATTITUDECONTROL
YAW ROLL PITCH
(POSITIVE) (POSITIVE) (POSITIVE)
(VIEW LOOKING EWD)
Cfl.14556.1
Figure 5-55
5-51
SectionVS-IVBStage
AUXILIARYPROPULSIONSYSTEM
QUANTITY INDICATOR
LANYARD ATTACH POINT
(TYPICAL FOR OXIDIZERAND FUEL
PROPELLANT TANK
ASSEMBLY
TANK
(N204)
STEEL
BELLOWS
DIZER
TANK
ULLAGE
PRESSURE
HELIUM TANK
TANK
ULLAGE
STEEL
BELLOWS
TANK
(MMH)
LINE FLANGES
(TYP)
X Ib>SCHE_T'C J
FLOW PATH
(TYPICAL FOR
BOTH UNITS)
FLANGE
GLASS ROVING--
FUEL FLOW ORIFICE--
VENT
VENT
ELECTRICAL
FUEL FLOW DIVIDER,
5-52
DOOR
I I;1!4GINE
i
NO.2
CESS
DOOR
CONNECT
POINTS (TYP)
GLASS ROVING
IMPREGNATEDSILICA
OVERWRAp
.MOLYBDENUM
3PRAYED
ZIRCONIACOATING
TRANSDUCER
)RIFICE <_PROPELLANT VALVES (4 FOR
FUEL, 4 FOR OXIDIZER) CH-14255.2
Figure 5-56
assembly first passes through a filter and then through two regula-
tors in series, both of which sense downstream pressure. During
normal operation, the primary regulator maintains regulated pres-
sure at 196 (+3, -6) psig. Should the primary regulator fail open,
the secondary regulator would maintain the output at 200 (+3,
-6) psig.
Ambient pressure sensing ports, provided on both regulators, fur-
nish the necessary ambient pressure references and provide a check-
out capability for enabling the regulated outlet pressure settings
to be checked. Regulator performance is evaluated by pressure
transducers installed immediately upstream and downstream from
the regulator. Regulated helium is fed through the quad check
valves to the ullage area of the fuel and oxidizer tanks.
Two sets of quadruple check valves are employed in the helium
pressurization subsystem; one set in the fuel tank pressurization
line and the other set in the oxidizer tank pressurization line. These
check valves are a redundant safety feature used to prevent contact
of fuel and oxidizer vapors in the pressurization subsystem should
both positive expulsion bellows develop a l_ak. Each set of check
valves consists of four check valves contained in one body connected
in a series-parallel arrangement. Failure of a check valve set re-
quires open failure of two check valves in series or closed failure
of two check valves in parallel.
The low pressure helium modules provide capabilities for ground
venting of propellant tank ullage as well as a means of establishing
pneumatic control of the expansion rate of the expulsion bellows
during loading and checkout. These modules also provide ground
and flight relief capability. The relief value cracking pressure is
325 to 350 psia; it reseats at 225 psia minimum. No command
venting capability during flight is provided.
Positive Expulsion Propellant Feed Subsystem. The positive expul-
sion propellant feed subsystem provides hypergolic propellant
transfer to the engines under zero and random gravity conditions.
This subsystem consists of an integral propellant tank assembly
containing two metal bellows for propellant expulsion, two propel-
lant control modules with propellant filters and auxiliary ports for
servicing operations and propellant manifolds for distribution of
propellants to the engines.
The propellant tank (figure 5-56) is an integral assembly consisting
of cylindrical compartments separated from each other by a helium
storage sphere. Each compartment contains a positive expulsion
bellows, with one bellows being used for fuel and the other foroxidizer. The bellows are constructed of welded stainless steel
convolutions incorporating a hemispherical movable end. A cor-
rugated liner is installed between the tank wails and the bellows
assembly. This liner serves as a guide for axial movement of the
bellows and as a bellow; vibration damper.
The bellows can be moved to either of its extreme positions (col-
lapsed or extended) through the application of a small differential
pressure across the bellows movable end. Pressurized helium, when
applied to the bellows movable end, compresses the bellows and
expels propellant on demand from the end of the tank. When the
bellows is completely collapsed, the movable hemispherical end
comes to rest against the mating tank bottom. This close stackingof the mating parts permits an expulsion efficiency of 98 percent.
The bellows assemblies have a minimum service life requirementof 200 cycles. One cycle consists of extension from the stacked
hieght (collapsed) to the fully expanded position and returned to
the stacked height.
A bellows position indicator connected externally to each propellant
tank, consists of a potentiometer connected to the bellows by means
of a flexible cable attached to a spring wound drum. The poten-
tiometer is connected to the drum and provides an indication of
bellows position. The position indicator assembly contains a mi-
croswitchwhichisactuated at 95 percent of the fully extended
bellows position (100 percent fill level). These microswitches per-
form no flight function and can be employed during servicing toindicate that the tank has been adequately loaded.
The propellant control modules filter propellants and provide access
into the propellant subsystem for servicing. The module contains
a propellant transfer valve, a recirculation valve, facility lines purge
check valves, and a system filter. The propellant transfer valve
is a direct operating, normally closed solenoid valve. This valve
is employed to fill and drain the attitude control system module.
The propellant recirculation valve is a direct acting, normally closed
solenoid valve with two independent poppets and seats. The two-
poppet design isolates the engine reeirculation line from the tank
recirculation line and all propellant flowing to the engine passes
through the filter. The filter has a 10-micron nominal and 25-micron
absolute rating.
The propellant manifold delivers propellant to the engine from
the propellant control module and provides a secondary path of
recirculation of propellant to the control module.
Engine and propellant control modules, which attach to the mani-
fold, use Marman conoseal flanges. Bellows are employed in the
lines to provide manifold flexibility.
Attitude Control Engines. Three 150-1bf TRW, Inc. engines (figure
5-56) are employed in each attitude control system module. The
combustion chamber is lined with ablatively refrasil material. The
engines have quadruple propellant or injector valves for redun-
dancy.
The thrust chamber is integrally fabricated and composed of three
major elements: the combustion chamber, the nozzle throat section,
and the nozzle expansion cone.
The engine has an expansion ratio of 33.9 to 1. The injector consists
of twelve pairs of unlike-on-unlike doublets arranged to minimize
hot spots in the combustion chamber. The valve side of the injector
is filled with a silver braze heat sink which reduces injector operat-ing temperature.
The engine was qualified for a total pulse operation of 300 sec.
During the 300-sec life requirement, the external wall temperature
does not exceed 600 ° F ..nd the maximum valve body external
temperature does not exceed 165 ° F. The maximum expected duty
cycle requirements on the Saturn S-IB/S-IVB is approximately 90
sec. The rocket engines have a reliability design goal of 0.9992
at a 90-percent confidence level while operating.
The engine propellant valve assembly consists of eight normally
closed, quad redundant propellant valves (four oxidizer and four
fuel), arranged in two series parallel arrangements. Dual failure
within the manifold fuel or oxidizer arrangement is required to
cause failure of the rocket engine assembly. An assembly closed
failure prevents any engine operation while an assembly open
failure results in continuous flow and loss of all propellant. Assem-
bly valve failure cannot occur unless two valves fail open in series
or two valves fail closed in parallel.
The injector valves provide positive on-off control of propellant
flow upon command from an external power source. Four valves,
integral in an assembly, are capable of simultaneous operation and
are synchronized to allow flow or terminate flow within 3 msec
of each other. The opening time for each valve assembly, defined
as the time from initiation of open signal to fully open valve
package, does not exceed 23 msec.
Each valve plunger and solenoid consists of a complete assembly
that may be removed from the propellant valve assembly to allow
thorough valve cleaning and solenoid replacement. Each valve
assembly (four fuel and four oxidizer valves) may be removed from
Section V S-IVB Stage
the rocket engine. All valve assemblies are interchangeable within
their own subsystem; oxidizer with oxidizer and fuel with fuel.
All external leak paths, at individual solenoid coils, are sealed by
Marman conoseals (excluding welded joints).
The leakage of the rocket engine fluids at each propellant injector
valve will not exceed a seepage rate of 0.03 in.3/24 hr during
exposure to ambient pressures and any propellant pressures up
to 185 psia.
APS Flight Operation.
The system on the S-IVB stage is enabled in flight after the S-IB
lower stage retrorockets have been ignited. First command actuation
occurs just prior to J-2 engine start to provide roll control of the
stage during engine burn. The single J-2 engine performs pitch
and yaw control only.
Helium supplied from a 3100-psia tank is regulated to 196 (+3,
-6) psig to expel propellants from each propellant tank to the
engines. During propellant flow, the liquid passes out of the bellows
tank into the engine mainfold for injection into the engines. See
figures 5-57 and 5-58.
Commands for control are provided by the instrument unit. Output
from a guidance platform indicating measured vehicle attitude is
received by the instrument unit and a comparison is made with
the desired or programmed attitude. If a deviation exists, the
instrument unit provides the required commands via the control
relay package to the attitude control system engine injector valves
for correction proportional to the magnitude of the deviation. The
engines operate in short pulse-type bursts and may operate 65 msec
or longer as required.
At J-2 engine cutoff the pitch and yaw control is activated and
all control (roll, pitch and yaw) remains active throughout the coast
phase.
Attitude Control System Performance.
The maximum impulse usage expected for SA-206 up to and
including CSM separation is 12.5 Ibm of oxidizer and 7.9 Ibm
of fuel per module. This includes S-IVB burn roll control, insertion
transients, maneuver to local horizontal and hold, and CSM separa-
tion transients. This consumption is 33% of the propellants loaded.
The total capacities are 39.4 Ibm of oxidizer and 23.6 Ibm of fuel
per module. The propellants remaining after CSM separation will
be used for passivation and the M415 experiment.
APS Servicing.
The North American Rockwell services supply the MMH and
N zO 4 to the APS modules through the McDonnell Douglas models
472 and 473 propellant loading carts, respectively. Two 0-100 psig
regulator gage assemblies, one fuel and one oxidizer, provide GN 2
for loading, pressurization, and purging of GSE and facilitylines.
The model 472 and 473 carts provide control for fuel and oxidizer
during propellant recirculate and fill procedures. APS servicing
is conducted while the mobile service structure is at the pad. All
servicing operations are accomplished through fittings on the aft
end of each APS module, except helium system purge and pres-
surization, which are accomplished through a single coupling on
the aft umbilical panel.
Figure 5-59 presents information pertinent to servicing the APS.
The APS propellant systems (figure 5-57) are purged with GN 2
prior to propellant loading. This purge clears the propellant filland return line and the fuel and oxidizer fill modules. After the
purge operation, ON 2 supplied through the fuel and oxidizer low
5-53
oectionV S-IVB Stage
(_ HIGH-PRESSURE TANK SUPPLIESHELIUM TO REGULATOR ASSY.
(_ REDUNDANT CHECK VALVES PREVENTHELIUM FLOW BACK TO THE UMBILICAL.
(_ PRIMARY REGULATOR REDUCESPRESS. TO 196 PSIG.
(_ BACK-UP REGULATOR FUNCTIONSONLY IF PRIMARy REGULATOR
FAILS.
(_ QUADRUPLE CHECK VALVES
PREVENT REVERSE FLOW FROM
TANK ASSYS INTO REGULATOR
ASSY.
(_) HELIUM PRESSURIZES FUEL AND
OXIDIZER LIt.LAGE (BELLOWS).
(_ RELIEF VALVES IN HELIUM LOW-
PRESSURE MODULES PREVENT
OVERPRESSURIZATION OF FUEL
AND OXIDIZER BELLOWS.
(_ INDICATORS PROVIDE CONTINUOUS
MONITORING OF FUEL AND
OXIDIZER QUANTITY.
(_ FUEL AND OXIDIZER MAINTAINED
IN SUPPLY MANIFOLDS, READYFOR INJECTION INTO ENGINES.
(_) IU FLIGHT CONTROL COMPUTER(FCC) COMMANDS OPEN THE FUEL
AND OXIDIZER CONTROL VALVES ON
ON ENGINE(S) REQUIRED FORATTITUDE CORRECTIONS.
(_) PRESSURIZED FUEL AND OXIDIZER
ENTER ENGINE COMBUSTION
CHAMBER AND IGNITE SPONTANE.
OUSLY. CONTROL VALVES CLOSE
UPON REMOVAL OF COMPUTERCOMMAND.
SECTION A-A
l_ll_ FUEL FILL
MODULE
,//_1 FILL & DRAIN, • u SOLENOID
SECTION B-B
SECTION C-C
SECTION D-D
SECTION E-E Q
150 LBF
ENG NO.
VD0078 j_
I_" VD_l--J- "E.
®Qo®•! ',_ ;_
PURGEL,NE__oj._'RECIRCULATIONLINE_IH "OXIDIZER FILL I H
AND DRAIN LINE nnluim_j
PURGE AND FUEL BELLOWS _]PRESSURIZATION LINE
LEGEND
m i m m OXIDIZER TANK PURGE
imam OXIDIZER (N204)
._ 3200 psi He
J 196 psi He
APS MODULE NO. I (UNIT 414)
APS MODULE NO. 2 (UNIT 415)
VXD0092 |
..... ENGINE VENT (FUEL) _ TELEMETRY INPUT
QUADRUPLE
CHECK VALVE
_:> C0|66 ASSY (OXIDIZER
OXIDIZER
O (N204) _ D0065
_:> VXN0039
VXN0040 •
INDICATOR •
O[_> VXN0037
TANK
lllllll•lll•I•lll•l•i
SOLENOID (NC)
VALVE
APS MODULE
NO. 2 SUPPLy
FILL MODULE
SOLENOID (NC)
RELIEF 325 [_ VXD009 I
TO 350 psld
CHECK VALVE
ASSY (FUEL)
FUEL I •I •I •
I •I •
I •
FILTER I •I •
I _FILTER | •
i \,.. •;URGE VEN
i _l_ SOLENOID I
! y (mc)_
I
I
i
BACK-UP REG
200 (+3, -6) psiOUTPUT
0
REG
196 (+3, -6) psiOUTPUT
[_> VXD0068
XD253
LINE
JLATION LINE
LINE
PURGE AND OXIDIZER BELLOWS
PRESSURIZATION LINE
CONSOLE
432A
AFT UMBILICAL
PLATE
LBF
ENG NO. 3
CH-14256-1
5-54Figure 5-57
Section V S-IVB Stage
OXIDIZERTANK
QUANTITY
OXIDIZER
TANK
HELIUM
RECIRCULATION
OXIDIZER
FILL AND
SUPPLY
L-----_.
r QUAD CHECK VALVES I
PRESSURE
TANK
FUEL TANK
QUANTITY
INDICATOR
APS TANK ASSY ]
/
IFUEL I
TANK jBELLOWS
t---BELLOWS•_ I _PURGE ANDI _ ! L RRESSUR,_T,O.
i "_::::_ PORTL I _ OVERmARD
i i J DUMP
I FUEL HELIUM I
QUAD CHECK VALVES 1
"I HELIUM PRESS. I
L REGOLATORASSY]
LOW PRESSUREMODULE
'--FUEL FILLAND
SUPPLY !
LINI
_--RECiRCULATION
LINE
RECIRCULATION
PORT_
J
OXI_PELL__T_I
CONTROL MODULE j
t-lgure 5-58
f_
j FUEL PROPELLANTCONTROL MODULE
RECIRCULATION
P°Z-7
1
PURGE PORT
I._-_ FILL- DRAIN
JJ PORT
I CH-1471g-1
pressure modules collapses the bellows to begin the propellantloading operations.
At the initiation of loading, a recirculation flow is first established
through each module. Propellants flow through the module propel-
lant lines and out through the recirculation valves. Flow is continued
until all trapped gas is eliminated by the recirculation procedure.
Once the lines are cleared of trapped gases, the recirculation valve
is closed and the bellows is allowed to expand by venting the GN 2
back through the low pressure helium modules allowing the tank
to be filled. Bellows position is monitored during loading by use
of position potentiometers installed on each tank assembly.
Helium usea tor propellant expulsion is loaded into the module
through a pneumatic service line connected to the stage throughthe stage umbilicals.
APS Fluids Summary.
Helium. Helium conforming to MSFC-SPEC-364A is used as a
pressurant. The purity of the helium shall be not less than 99.95
percent helium by volume when tested in accordance with MSFC
requirements. Impurities shall not exceed: oxygen, 90 ppm by
volume: hydrogen, no test required; nitrogen, 360 ppm by volume;inert gases, no test required.
5-55
Section V S-IVB Stage
FUNCTIONAL
OPERATION MEDIA
.Ii lllllllJWJ b|llllL |! BILIIILI/.IE _14.•111l_141.l| Ill,
FLOW TIMEPRESS. TEMP. RATE REQUIRED QUANTITY
(pslg) (OF) (GPM) (MIN) (L8)
REMARKS
FUEL LOADING MMH 25 TO 35 70 TO 90
OXIDIZER LOADING N20 4 25 TO 35 70 TO 90
FUEL CIRCULATION
(/=RELOADING) MMH 25 TO 35 70 TO 90
OXIDIZER CIRCULATION N20 4 25 TO 35 70 TO 90(PRELO ADING)
OXIDIZER AND FUEL
LOADING AND TRANSFER
LINES PURGE
(POSTLOADING)
GN 2 25 TO 35 AMB
OXIDIZER TANK BELLOWS
AND FUEL TANK BELLOWS GN2 35 TO 45 AMB
COLLAPSING PRESSURE
0.STO 2O 7.5
0.5TO 2.5 7.5
1.5TO 2.5 15
1.5 TO 2.5 15
1.0 TO 2.0 15
STATIC N/A
STATIC N/A
23.6
39.4
REQUIREMENTS ARE GIVEN FOR
ONE MODULE. LOADING ACCU-
RACY IS DEPENDENT ON MISSION
REQUIREMENTS. ACCURACy CAN
BE HELD WITHIN + I% OF MASSREQUIRED.
109 PRIOR TO FILLING, FUEL AND
OXID ARE CIRCULATED THROUGH
SYSTEM WITH BELLOWS COLLAPSED.179
THE PURGE IS TO INERT THE LINES.
REQUIREMENTS FOR ONE MODULE,
TWO REQUIRED.
REQUIREMENTS FOR ONE MODULE,0.015 NOM TWO REQUIRED. NOT A MAIN
UMBILICAL.
OXIDIZER AND FUEL GN 2 50 + 5 AMBULLAGE PRESSURI ZATION
N/A
0.160HELIUM STORAGE SPHERE HELIUM 3100 + 100 AdVIB 0.75 LB/MIN MAX 65 + 5
Figure 5-59
CH.14405.1
Nitrogen. Nitrogen conforming to MSFC-SPEC-234A is used for
purging and pressurization during loading. The nitrogen purity shallnot be less than 99.983 percent nitrogen by volume when tested
in accordance with MSFC requirements. Impurities shall not ex-
ceed: oxygen, 150 ppm by volume; carbon dioxide, no test required;
inert gases, no test required.
Freon 113 (TF 113). Freon 113 is used to flush the oxidizer system
in accordance with MSFC requirements. The product requirementfor TF 113 is the same as for Freon 12.
Nitrogen Tetroxide (Inhibited). The nitrogen tetroxide (N zO .0 per
specification MSC-PPD-2B is used as an oxidizer propellant. It
contains 0.8 -+- 0.2% by weight nitricoxide as an inhibitor. See
figures 5-60 and 5-61 for N 20 4 properties and vapor pressure.
Monomethylhydrazine. The monomethylhydrazine (MMH) per
specification MIL-P-27404 is used as a fuel propellant for the fuel
system. See figures 5-62 and 5-63 for MMH properties and vapor
pressures.
APS Measuring.
• Thirty-eight measurements, nineteen per module, are taken during
flight and telemetered to ground receiving stations (figure 5-64).
All measurements are recorded for postflight evaluation. The flightcontrol measurements are monitored in real-time at mission control
center. The results of the flight control measurements are used
in decision making affecting the mission in contingency situations.
ELECTRICAL.
The stage electrical system produces and distributes all ac and dc
power for flight functions. Four batteries, two each in the forward
and aft skirts, supply 28-Vdc and 56-Vdc primary power. This power
buses directly to an associated distribution assembly. From here
power routes through networks operational equipment and secon-
dary power sources (to convert and regulate for specialized func-
tions). There are four primary networks-one from each battery.
Two networks I +4D21 and +4D31 originate in the forward skirt.
5-56
dary power sources (to convert, invert and regulate for specialized
functions). There are four primary networks-one from each battery.
Two networks (+ 4D21 and + 4D31) originate in the forward skirt
PROPEPTIES OF N20,
COMMON NAME:
NITROGEN TETROXIDE, NITROGEN PEROXIDE OR LIQUID
NITROGEN DIOXIDE.
CHEMICAL FORMULA:
N2 0 4 2NO 2
MOLECULAR WEIGHT:
92.016
PHYSICAL PROPERTIES:
FREEZING POINT .................... 1 ] .84°F
BOILING POINT ...................... 70.07°F
CRITICAL TEMPERATURE ............ 316.8OF
CRITICAL PRESSURE .................. t469 PSIA
LIQUID DENSITY (77°F) ............. 11.94 LBS/GALAPPEARANCE .......................... AT ROOM TEMPERATURE
A HEAVY BROWN LIQUID.
AS TEMPERATURE DECREASES,
COLOR BECOMES LIGHTER.
FUMES ARE YELLOWISH TO
REDDISH BROWN.
ODOR ................................... CHARACTERISTIC PUNGENT
ODOR
MAXIMUM ALLOWABLE
CONCENTRATION ................... 2.5 PPM
CHEMICAL PROPERTIES:
A CORROSIVE OXIDIZING AGENT. HYPERGOLIC WITH UDMH, HYDRA-
ZINE, ANILINE, FURFURYL ALCOHOL AND SOME OTHER FUELS.
NOT SENSITIVE TO HEAT, MECHANICAL SHOCK OR DETONATION.
IS NONFLAMMABLE WITH AIR, HOWEVER, IT CAN SUPPORT COMBUSTION
WITH COMBUSTIBLE MATERIALS. VERY STABLE AT ROOM TEMPERATURE.
Figure 5-60
C.H 14539
Section V S-IVB Stage
NITROGEN TETROXIDE
VAPOR PRESSUREPROPERTIES OF MMH
130"
120-
110-
100-
90-
80-
70-
60-
50-
40-
PSIA
30- / o 1t.84 3.0 FREEZING POINT
i
20- ._ • 70.07 14.7 _OILING POINT
10- _ / 316.8 1469.0 CRITICAL POINT
0 I I I I I I I I I I I I I I I I I I I
0 20 40 60 80 100 120 140 160 180
DEGREES FAHRENHEIT (°F) C.H 14535
Figure 5-61
and two others (+4DII and +4D41) originate in the aft skirt.
All networks supply 28 Vdc, except +4D41 which supplies 56 Vdc.
Each network is independent of the others. Figure 5-65 shows the
distribution network arrangement.
DESIGN GROUND RULES.
The electrical system design is analogous to those of earlier weapon
systems (Thor) and space vehicles (S-IV stage). These proven con-
cepts influence the electrical system ground rules.
Power and Signal Distribution.
The stage has two separate primary power distribution networks
and source power completely independent from other vehicle stages.Each network has switching provisions to selectively alternate be-
tween ground source power (prelaunch) and stage battery power
(flight). The stage wiring arrangement limits voltage drop to a 2-V
(maximum) level from bus to load. Placing separate networks above
and below the propellant tanks reduces cabling and eliminates need
for excessive power lines in the connecting cable tunnels (see Struc-
ture). Electrical bonding, techniques (metal-to-metal) ensure a uni-
potential structure. Electrical circuits are insulated from their re-
spective cases. These cases bond to the structure for personnel safety
and short circuit protection. Low current signals (measurements
and telemetry) use separate networks; the cables are shielded and
spaced to prevent electrical interference with high current lines.
Packaging and Installation.
Electrical components (or assemblies) are modular wherever possi-
ble. This packaging concept increases reliability, speeds part chan-
geout, and reduces the number of cable interconnections. The
modules, encapsulated when design permits, bond to open assembly
panels for convenient maintenance and replacement. Physical para-
meters and heat dissipation requirements sometimes prevent en-
capsulation (epoxy) of equipment; instead, these components areinstalled in pressurized boxes. Mounting each assembly on its center
of gravity minimizes induced vibration. Most electronic equipmentis in the forward skirt where it mounts on environmental (cold-plate)
COMMON NAME:
MONOMETHYL HYDRAZINE (MMH)
CHEMICAL FORMULA:
CH3NHNH 2
MOLECULAR WEIGHT:
46.075
PHYSICAL PROPERTIES:
FREEZING POINT ........................... -62.5°F
BOILING POINT ............................. 192.5 °F
CRITICAL TEMPERATURE .................... 593.6 OF
CRITICAL PRESSURE ......................... 1195.1 PSIA
LIQUID DENSITY (77°F) .................... 7.29 LBS/GAL
APPEARANCE ................................. CLEAR, OILY WATER
WHITE LIQUID.
ODOR ........................................... AMMONIACAL
MAXIMUM ALLOWABLE
CONCENTRATION .......................... 0.5 PPMFLAMMABILITY LIMITS ..................... 4 TO 97% IN AIR
CHEMICAL PROPERTIES:
MMH IS NOT SENSITIVE TO IMPACT OR FRICTION, HOWEVER, ITS VAPOR
WILL PROPAGATE DETONATION WITHIN THE FLAMMABILITY LIMITS.
EXTREMELY MISCIBLE IN WATER.
WILL REACT WITH CARBON DIOXIDE AND OXYGEN IN AIR.
IRON, COPPER, MOLYBDENUM, AND THEIR OXIDES CAUSE MMH TO
DECOMPOSE.
VERY CAUSTIC .
VERY HYGROSCOPIC.
IGNITES SPONTANEOUSLY IF EXPOSED TO A LARGE SURFACE (SUCH AS
SATURATED RAGS).
FREEZING HAS NO EFFECTS ON THE CHEMICAL PROPERTIES.
Figure 5-62
C-H 14534
MMH VAPOR 'RESSURE
375 -
350-
325-
300-
275-
250-
225-
200-
175-
150-
125-
100-
75-
50-
25-
o T,
TEMP. OF PSIA
0.0 FREEZING POINT
,// _92.5 14.7 BOILING POINT
' I , 593,6
1195.1 CRITICAL POINT
I I i i I100 200 300 400 500 600
DEGREES FAHRENHEIT (OF) C,H 14536
Figure 5-63
5-57
SectionVS-IVB Stage
panels; the aft skirt has only ambient panels. Figures 5-2 and 5-3show the equipment location in both skirt areas.
BATTERIES.
Studies proved that the best type of power source for the stage
mission is a battery. Consideration was given to various other
sources such as fuel cells, sterling cycle engines, and solar arrays.
However, batteries best satisfy the power level requirements for
the mission time duration. Weight and voltage regulation charac-
teistics are basic reasons for using primary (one-shot) batteries
instead of secondary (rechargeable) batteries. The original systemdesign required only two batteries: one in the forward skirt for
noise-free power to the data acquisition, propellant utilization, and
range safety systems; another in the aft skirt to provide powerswitching and sequencing functions. Subsequent requirements for
redundant range safety and emergency detection systems forced
the addition of a second battery in the forward skirt. High power
requirements of the chiUdown pump motors and auxiliary hydraulic
pump motor forced the addition of a second battery in the aft
skirt. A 56-Vdc battery was selected because it drives these motors
more efficiently and reduces complexity in the chilldown invertercircuitry.
Physical Construction,
The same manufacturer (Eagle Picher) produces all four stage
batteries; in fact, all vehicle batteries. Section IV gives a detailed
description of battery construction that is basically true for these
batteries. Some features change such as mounting provision, elec-
trical connection, overall size, cell configuration, and rated capacity.
Notice that forward battery no. I is packaged as two separate units.
Electrical Characteristics.
The four batteries use silver oxide (pos) and zinc (ueg) electrodes
in a potassium hydroxide electrolytic solution to produce the elec-
tromotive force. Forward batteries no. l and no. 2 and aft battery
no. I have output voltages of 28 ± 2 Vdc; aft battery no. 2 has
an output of 56 -+- 4 Vdc. These outputs hold true when each
battery loads to the profile in figure 5-66.
The following list gives the nominal capacity and the predictedusage for each battery:
a. Fwd No. l Battery: 300 A-hr capacity, 40 percent usage.
b. Fwd No. 2 Battery: 5 A-hr capacity, 76 percent usage.
c. Aft No. l Battery: 58 A-hr capacity, 87 percent usage.
d. Aft No. 2 Battery: 22 A-hr capacity, 51 percent usage.
Internal heaters maintain battery temperature close to the normal
operating range of 70 to 90 ° F. Circuitry inside the battery provides
inflight monitoring of battery voltage (bridge network), current(current transformer), and temperature (thermistor). These mea-
surements are listed in figure 5-67 with the ground station thatdisplays them.
CHILLDOWN INVERTERS.
Motor-driven pumps, one each for the lox and fuel systems, circulate
the liquid propellants to remove latent heat and stabilize temper-
ature at the J-2 engine turbopump inlets before engine start (see
Propulsion). These two pump motors operate on ac voltage from
separate power supplies. These secondary power supplies, called
chilldown inverters, receive 56 Vdc from aft battery no. 2 (+4D41
bus) and invert it to 3-phase current with 1500 VA capacity. The
lox inverter is operable 5 min before liftoff; the fuel inverter powers
5-58
NUMBER
XC166-414
XC167-415
*XC0168-414
*XC0169-415
*XC0170-414
*XC0171-415
*D0063-414
*VXDO064-414
* D0065-414
"130066-415
* D0067-415
*VXDO068-415
*[30069-415
*D0084-414
*VD0078-414
*VD0079-414
* VD0080-414
*VD0081-415
*VD0082-415
*VDO083-41S
VXD0089-414
VX D0090-414
VX00091-415
VXD0092-415
VXD0093-414
VXD0094-414
VXD0095-415
VXD0096-415
*XD252-414
*XD253-415
VK 132-404
VK133-404
VK 134-404
VK 135-404
*VXN0037-414
*VXN0038-415
*VXN0039-41S
*VX N0040-415
NAME
TEMP, He SPHERE MODULE 1
TEMP, He SPHERE MODULE 2
TEMP, OXIDIZER TANK OUTLET MODULE I
TEMP, OXIDIZER TANK OUTLET MODULE 2
TEMP, FUELTANK OUTLET MODULE 1
TEMP, FUELTANK OUTLET MODULE 2
PRESS, FUEL SUPPLY MANIFOLD MODULE 1
PRESS, He REGULATOR INLET MODULE 1
PRESS, He REGULATOR OUTLET MODULE 1
PRESS, OXIDIZER SUPPLY MANIFOLD
MODULE 2
PRESS, FUEL SUPPLY MANIFOLD MODULE 2
PRESS, He REGULATOR INLET MODULE 2
PRESS, He REGULATOR OUTLET MODULE 2
PRESS, OXIDIZER SUPPLY MANIFOLD
MODULE I
PRESS, ATTITUDE CONTROL CHAMBER 1-I
PRESS, ATTITUDE CONTROL CHAMBER I-2
PRESS, ATTITUDE CONTROL CHAMBER 1-3
PRESS, ATTITUDE CONTROL CHAMBER 2-I
PRESS, ATTITUDE CONTROL CHAMBER 2-2
PRESS, ATTITUDE CONTROL CHAMBER 2-3
PRESS, FUELTANK ULLAGE MODULE 1
PRESS, OXIDIZER TANK ULLAGE MODULE 1
PRESS, FUELTANK ULLAGE MODULE 2
PRESS, OXIDIZER TANK ULLAGE MODULE 2
PRESS, FUEL TANK OUTLET MODULE 2
PRESS, OXIDIZER TANK MODULE 1
PRESS, OXIDIZER TANK OUTLET MODULE 2
PRESS, FUELTANK OUTLET MODULE 2
PRESS, He PEG IN MODULE 1
PRESS, He REG IN MODULE 2
EVENT, ENGINE 1-I/1-3 FD VLV OPEN
EVENT, ENGINE 1-2 FD VLV OPEN
EVENT, ENGINE 2-1/1-3 FD VLV OPEN
EVENT, ENGINE 2-2 FD VLV OPEN
QTY, OXIDIZER TANK MODULE 1
QTY, OXIDIZER TANK MODULE 2
QTY, FUEL TANK MODULE I
QTY, FUELTANK /'._DDULE 2
RANGE
3600 TO 660°R
360 ° TO 660°R
460 ° TO 590°R
460 ° TO 590°R
460 ° TO 590°R
46Oo TO 590°R
0 TO 400 PSIA
0 TO 3500 PSIA
0 TO 400 PSIA
0 TO 400 PSIA
0 TO 400 PSIA
0 TO 3500.PSIA
0 TO 400 PSIA
0 TO 400 PSIA
0 TO 200 PSIA
0 TO 200 PSIA
0 TO 200 PSIA
0 TO 200 PSIA
0 TO 200 PSIA
0 TO 200 PSIA
0 TO 400 PSIA
0 TO 400 PSIA
0 TO 400 PSIA
0 TO 400 PSIA
0 TO 400 PSIA
0 TO 400 PSIA
0 TO 400 PSIA
0TO 400 PSIA
0 TO 3500 PSIA
0 TO 3500 PSIA
0 TO't0 IN.
0 TO 10 !N.
0 TO 10 IN.
0 TO 10 IN,
MEASUREMENT NUMBER PRERXES INDICATE THE FOLLOWING:
*FLIGHT CONTROL; V-ESE DISPLAY; X-AUXILIARY DISPLAY
CH-14386.
Figure 5-64
up 23 sec later. See Sequencing for power off commands in flight.
The two inverters are identical units developed for this specific
task. Previous converters (using 28 Vdc) were 80 percent heavier,
had 80 percent more components, and were 30 percent less efficient.
Section V S-IVB Stag_
TO"
FWD BATTERY COMMAND
HEATERS PWR TRANSFER ESE ESE
FORWARD
SYSTEM
AFT
SYSTEM
ESE
pWR
COMMAND TO
PWR TRANSFER AFT BATTERY
EXT INT HEATERS
ON IGN
PWR PWR
JR ENG
ELEC CONT
PKG d01A4
COMMAND
POWER
ESE TRANSFER
pWR EXT INT
Notes " *
:_ TYPICAL ON SIGNAL
CONDITIONING RACKS
CIRCLED NUMBERS INDICATE
T,MES(_) 7 HR 15 MIN BEFORE LIFTOFF( 2 ) I HR 20 M,N BEFORE LIFTOFF
(_ 9 MIN 30 SEC BEFORE LIFTOFF _/- + •
(_ 10 MIN 30 SEC BEFORE LIFTOFF _"_": _,: _
9 MIN 50 SEC BEFORE LIFTOFF _ '_ "
_) 9 MIN 56 SEC BEFORE LIFTOFF
(_ 50 SEC BEFORE LIFTOFF
(8) 13 HR 22 MIN BEFORE LIFtC)FF
Figure 5-65
AUX HYDRAULIC
PUMP MOTOR
i 403BI
LOX
CHILLDOWN
INVERTER
404A74AI
FU
CHILLDOWN
INVERTER 404A74A2 CH.14381-1
5-59
SectionVS-IVB Stage
PU STATIC INVERTER-CONVERTER.
This unit is a secondary power supply that regulates voltages tothe PU electronics assembly (see Propulsion) for various PU func-
tions. The unit is operable 5 hr 45 min before liftoff; a switch
selector function switches power off in flight (see Sequencing).
Forward battery no. 2 (+4D21 bus) provides 28-Vdc operating
power and the unit supplies the following specialized outputs:
a. 115 Vac, 400 Hz drives the bridge rebalancing servo motors.
b. 21 Vdc operates PU bridge networks.
c. 5 Vdc excites PU fine and coarse mass potentiometers.
EXCITATION MODULES.
Two different types of signal excitation modules, 5 Vdc and 20Vdc, provide excitation for transducers and instrumentation ne-
tworks. As such, these excitation modules are secondary power
supplies for the instrumentation system. Ground rules specify that
measurements must be in the 0 to 5 Vdc range when applied to
the digital data acquisition system. Since measurement signals
originate in various forms, many require 0 to 5 Vdc conditioning.
Several types of solid-state conditioning modules perform this
function; for example, dc amplifiers, temperature bridge networks,
and frequency-to-de converters. These conditioning modules receive
-20 Vdc, +20 Vdc and +28 Vdc operating power. In addition,
the measuring networks use a regulated 5-Vdc reference for calibra-
tion at ground receiving stations. Since 28 Vdc is available from
the primary netowrks, secondary power sources supply the other
converted voltages for measurements excitation.
5-Vdc Excitation Module.
This module contains solid-state electronic circuitry that transforms
an unregulated 28-Vdc input voltage into accurately regulated
output voltages. It consists of two conversion sections: one section
supplies up to 400 mA at 5 Vdc and up to 100 mA at -20 Vdc;
the other section supplies a 10-Vpp, 2000-Hz square wave and
has a capacity of 5 VA. The 5-Vdc output is monitored as a flight
measurement (see figure 5-67). Three modules are on the stage;
two are in the forward skirt, and one is in the aft skirt (see figures
5-2 and 5-3).
20-Vdc Excitation Module.
This is a conventional de-to-de converter with regulation circuitry.
It receives unregulated 24 to 30 Vdc and regulates the output to
20 ± 0.04 Vdc. A module mounts on all fbur signal conditioning
racks (one in the forward skirt and three in the aft skirt) to power
measuring system components.
DISTRIBUTORS.
There are five distribution assemblies that bus battery power to
PREDICTEDBATTERYLOADPROFILES
5-60
8O
70
68.2A
PROFILE KEY l
EVENT LOAD DURATION POWER SOURCE/CODE
FUEL CHILLDOWN INVERTER I TO 2 AFT BAT. NO. 2 i
LOX CHILLDOWN INVERTER 1 TO 3 AFT BAT. NO. 2
AUX HYDRAULIC pUMP 1 TO 6 AFT BAT. NO. 2
ENGINE IGNITION 3 TO 4 AFT BAT. NO. I lira===
APS PULSED (10A SPIKES) S TO 7 AFT BAT. NO. I ==n==mRS CMD SYS AND TM TRANSMITTERS 1 TO 8 FWD BAT. NO. I ............
RS CMD SYS AND INVERTER-CONVERTER I TO 9 FWD BAT, NO. 2 .=u..,m=
40.2A
II
0 + 1:40 3:20
22,2A
4.2A
5:.00 6:.4O
TIME AFTER LIFTOFF (MIN:SEC)
Figure 5-66
ENGINE
CUTOFF,
L10:00 11:40
CH-14506.2
NAME RANGE
*C102-411 TEMP,FWDBATTERYNO. 1, UNIT 1 460TO660°R
*C103-411 TEMP,FWDBATTERYNO. 2 460TO660°R
*XC104-404 TEMP,AFTBATTERYNO. 1 460TO660°R
*XC105-404 TEMP,AFTBATTERYNO. 2 460TO660°R
*C211-411 TEMP,FWDBATTERYNO. 1, UNIT2 460TO660°R
*XM14-404 VOLT,AFTBATTERYNO. 1OUTPUT 0TO 40V
*XM15-404 VOLT,AFTBATTERYNO. 2 OUTPUT 0TO 80V
*XM16-411 VOLT, FWDBATTERYNO. 1OUTPUT 0TO40V
*XM18-411 VOLT, FWDBATTERYNO. 2OUTPUT 0TO40V
*XM19-411 CURRENT,FWDBATTERYNO. 1LOAD 0TO200A
*XM20.-411 CURRENT,FWDBATTERYNO. 2 LOAD 0TO20A
*XM21-404 CURRENT,AFTBATTERYNO. 1 0TO200A
"XM22-404 CURRENT,AFTBATTERYNO. 2LOAD 0TO200A
*VM24.-411 VOLT,5.-VEXCITATION/COD,FWDI 4.5TO 5.5V
*VM25-404 VOLT,5-V EXCITATIONMUD, AFT 4.5TO 5.5V
*VM68-411 VOLT,5-V EXCITATION/COD,FWD2 4.5 TO5.5 V
MEASUREMENTNUMBERPREFIXESINDICATETHEFOLLOWING:*FLIGHTCONTROL;V-ESEDISPLAY;X-AUXILIARYDISPLAY
CH-_I_
Figure 5-67
the forward skirt distributes primary power from both batteries
in that area. Since 56-Vdc and 28-Vdc batteries are in the aft skirt,
separate distributors accommodate them. Two other distribution
assemblies, called control distributors, provide no flight functionand are not shown in the distribution diagram; they provide ESEinterface via the umbilicals for control and talkback functions
during prelaunch operations. Notice that the power distribution
assemblies use modular construction as mentioned in the designground rules. For example, each motorized transfer switch between
battery and primary bus is contained in separate modules. An
identical module is on each distribution assembly. There are alsorelay modules, bus modules, and connector modules that house
the complete assembly circuitry.
SEQUENCING.
As on all stages, a switch selector is the communications link
between the stage and LVDC where the flight sequencing program
is stored. Details on switch selector operations are outlined in
Section IV, Sequencing. As shown in figure 5-68 the switch selector
issues its commands to the sequencer assembly. The sequencer
assembly is modular in construction and contains logic circuitry
to perform its functions. It has magnetic latching relays, non-latch-
ing relays, transistor networks, and diodes that enable or disable
circuits, or switch power on-and-off to perform the various switchselector commands. Notice that some commands are repetitious,
such as tank venting during the coast mode which is explained
in AS-206 S-IVB Orbital Sating subsection.
INSTRUMENTATION SYSTEMS.
The S-IVB stage instrumentation systems consist of measuring
subsystems and one PCM/DDAS telemetry link. The major in-
Section V S-IVB Stage
strumentation equipment for the S-IVB stage is located on theinterior of the forward and aft skirt assemblies.
MEASURING SYSTEM.
The S-IVB measuring system consists of various types of trans-
ducers, four signal conditioning rack assemblies, and a central
decoder assembly and associated channel decoder assemblies. Ap-
proximately 252 measurements made on the S-IVB stage of the
AS-206 vehicle provide both preflight and inflight data. The mea-
surement coding system employed in the S-IVB stage is the same
as that used on the S-IB stage and has been previously describedin Section IV of this manual. Those measurements that are unsuita-
ble for use by the telemetry system in original form are modified
by signal conditioning modules before being routed to the telemetry
multiplexers. These signal conditioning modules mount on four
signal conditioning racks; three in the aft skirt and one in the
forward skirt. A central decoder assembly and associated channeldecoder assemblies translate coded commands to calibrate the
measuring system. The remote automatic calibration system
(RACS) permits a remote calibration of the measuring system prior
to launch. During vehicle checkout, calibration of measurements
is accomplished through the RACS and various corrections can
be made. Prior to launch, the RACS is operated to determine if
the system drifts or deviates from the final adjustments. The data
obtained is used to correct the flight data for more accurate mea-surements.
TELEMETRY SYSTEM.
The S-1VB telemetry system (figure 5-69) is a PCM/DDAS (pulse-
code modulated/digital data acquisition system) link (CPI) that
transmits real-time checkout data before launch and measuring
program information during flight. The system consists of two
Model 270 multiplexers, a remote digital submultiplexer, a re-
mote analog submultiplexer, a Model 301 PCM/DDAS assembly,
and an RF assembly. Components of the telemetry, system arelocated in the forward skirt, with the exception of the multiplexersthat are located in the aft skirt. The use of two time division
multiplexers, a digital submultiplexer, and an analog submulti-
plexer allows increased data transmission capability. The multi-
plexers sample measurement inputs to produce a train of varying
amplitude output pulses. The PAM (pulse-amplitude modulated)data from the multiplexers are converted to digital words in the
Model 301 PCM/DDAS assembly, which then encodes the digitalwords into a data frame. The data frame is transferred to the RF
assembly where it modulates the RF carrier. The RF signal is
transmitted through two antennas installed on opposite sides of
the vehicle by means of a bidirectional coupler and an RF powerdivider. A coaxial switch transfers the telemetry signal to RF
dummy load during checkout. Two power detectors measure the
forward and reflected power from the directional coupler. These
measurements are telemetered to ground stations and are used to
determine the VSWR and the actual RF power transmitted from
the vehicle. The PCM/DDAS feeds data through a coaxial cable
to the DDAS ground receiving station during prelaunch checkout.
With the exception of the remote analog submultiplexer (RASM),the telemetry components in the S-IVB stage are common to the
S-IB stage and are discussed in Section IV. The RASM is a PAM
analog submultiplexer, which provides the capability to remotely
locate any 6 primary channels of a -Model 270 time division
multiplexer. The RASM has 6 data output channels, each of whichhas a capacity of 10 submultiplexer channels. Each submultiplexed
channel is sampled 12 times per second.
Approximately 178 measurements originating in the S-IVB stage
are routed through multiplexer DPIB0 to the IU PCM/DDAS
assembly for transmission of these 178 measurements; approxi-
mately 78 are also routed through multiplexer CPIB0 for transmis-
5-61
tionV S-IVB Stage
ELECTRICALSECUENCINGTHIS CIRCUITRY IS SIMPLIFIED TO SHOW ONLY THE RELAYS AND DIODES NECESSARY TO EX-
PLAIN THE RESULTS OF THE SWITCH SELECTOR COMMANDS. THE TABLE LISTS SEQUENCED
EVENTS UNTIL CSM SEPARATION. CERTAIN EVENTS CONCERNING TM CALIBRATION AND
TANK VENTING CONTINUE UNTIL END OF S-IVB ACTIVE LIFETIME. THIS SEQUENCE
DOES NOT INCLUDE STAGE PASSIVATION OPERATIONS (SEE ORBITAL SARNG).
J_ S-IB PROP LEVEL SENSOR ACTUATION
S-IB O.B. ENG CUTOFF
J_" S-IB/S-IVB SEPARATION
j_ TURNED ON APPROX 10 MIN 30 SEC
BEFORE LIETOFF ON SW SEL CHAN 28
TURNED ON APPROX 9 MI N 30 SEC
BEFORE LIETOFF ON SW SEL CHAN 22
J_ TURNED ON APPROX 9 MIN 56 SEC
BEFORE LIFTOFF ON SW SEL CHAN 58
OPEN APPROX 3 SEC BEFORE L[FTOFF
ON SW SEL CHAN 80
NOT USED IN FLIGHT
[_ TURNED ON APPROX 7 HR 15 MINBEFORE LIFTOFF ON SW SEL CHAN 7
ONLY ONE OF THE THREE ULLAGE
/
RO KETS,SSHOWNPORCLAR,TyTHE/"'OTHER TWO CHARGE AND FIRE THROUGHPARALLEL FUNCTION OF SAME RELAYS
[_ THE FLIGHT PROGRAM P,,_ANS FOR
A GUIDANCE CUTOFF _ BASED
ON ADEQUATE VELOCITY. "_/,_;, I,_'_COMMAND (_) STARTS T4 AT 9 MIN ','_s0.1 SECAFTERurrOEF. HOWEVERPROPELLANT DEPLE'nONCUTOFF(_IS ARMED TO PROTECT AGAINST
PROPELLANT STARVATION. . .,_l_ _
HELD OPEN UNTIL PHYSICAL
SEPARATION OF S-IB/S-IVB
STAGES AT APPROX 3. I SEC
AFTER S-IB LEVEL SENSOR
ACTUATION
_. OPENEDBYMA,NSTAGE _THRUSTOKS;GNA_ _!_TURNED ON BY LOX TANK PRESS
SOL VLV CLOSE CMD & LOX TANK
PRE-PRESS FLT CONTL PRESS SW _SWITCH _
DEENERGIZED _.SELECTOR
I_" S-IVB ENGINE CUTOFF
JNCF
"IICLI
;ITAI
MPU
A28
UNC
HICL
_TA
)APTI
3A29
iAUX HYD
5-62
m
Figure 5-68 (Sheet 1 of 2)
SWITCH
SELECTOR
CHAN
79
62
It>54
83
80
103
13
49
5
34
56
10
59
23
9
27
11
68
6
35
55
57
16
88
73
50
5
34
97
12
48
12
48
42
99
82
9t
79
104
98
6
35
29
37
95
96
8
100
77
78
TIME
BASE
(SEC)
T 1 +6.0
T I + 127.5
T I + 128.5
T2 + 0.0
T2 ÷ 3.6
T2 + 4.3
T3 ÷ 0.0
T3 + 0.2
T3 + 0.3
T 3 + 0.4
T3 +0.5
T3 +0.8
T 3 + 0.9
T3+ 1.1
T3 + 1.3
T3 + t.9
T3+2.1
T3 + 2.3
T3 + 2,7
T3 +3.2
T3 + 3.7
T3 + 5.3
T3 + 8.7
T3 + 8.9
T3 + 10,2
T 3 + 13.3
T 3 + 13.7
T 3 + 19.3
T3 + 19,5
T3 + 24.0
T3 + 302.9
T3 + 328.1
T 3 + 328.3
T3 + 400
T 4 - 0.2
T4-0,1
T4 + 0.0
T4+0.1
T4+0.2
T4+0.3
T4+0,4
T4+0.5
T4÷0.6
T4+0.8
T4 + 1.0
T4 + 1.8
T4+2.2
T4+2.4
T4 + 3.9
T4 + 200.0
T4 + 203,0
T4 + 205.0
T4 + 240.0
T4 + 320.0
T4 + 323.0
T4 + 325.0
CH.20130
Section V S-IVB Stage
PUMP
SEQUENCER 404A3
U/R IGNITIOi"
404A3
Figure 5-68 (Sheet 2 of 2)
CH-14313-I
5-63
Section V S-IVB Stage
sion from the S-IVB stage. Approximately 73 measurements origin-
ating outside the S-IVB are routed through multiplexer CPIA0,
located in the IU, to the S-IVB stage PCM/DDAS assembly for
transmission. These measurements originate in the spacecraft, IU,
and S-IB stage. This configuration is necessary for the measuring
program to satisfy battery life requirements, EDS requirements,
and requirements for redundancy of other flight critical measure-ments.
ENVIRONMENTAL CONDITIONING.
FORWARD SKIRT.
The S-IVB stage forward skirt area is conditioned from a ground
environmental control system during preflight preparations through
a purge manifold located in the instrument unit (see Section VI).
The system provides ventilating air to the compartment during
the preflight phases of IU checkout. During LH 2 loading operations
the system maintains an inert GN 2 atmosphere in the compartment
to preclude accumulation of combustibles. No inflight purge is
required.
During flight, a cold-plate system provides thermal conditioning
for forward skirt equipment. A maximum of 16 cold plates connect
to the IU thermal conditioning system through a parallel hookup
of supply and return manifolds. SA-206 contains five cold plates
and 11 cold plate simulators that maintain normal system flow
characteristics through calibrated orifices where panels are not
needed. The cold plates serve as mounting surfaces and heat ex-
changers for heat-generating electronic equipment. A silicate ester
(Oronite FIo-Cool 100 dielectric coolant) liquid flows from the IU
thermal conditioning system through the cold plate supply mani-
fold, through the series-flow heat exchangers, into the return mani-
fold and back to the IU thermal system for conditioning. A fluid
minimum flowrate of 7.4 gpm and maximum temperature of 70 °
F is required to maintain the cold plates at a maximum design
temperature of 80 ° F for acceptable component heat extraction.
Radiation shields are installed over electronic components and
attached to the mounting panels to assist in maintaining acceptable
cold plate and equipment temperature. The thermal conditioning
system is operational during prelaunch electronic componentoperations and during flight.
AFT SKIRT.
The aft skirt thermal conditioning system provides for thermal
conditioning of launch critical electronic equipment, the attitudecontrol system, and the hydraulic accumulator-reservoir while
purging the aft skirt and interstage area prior to launch. Design
requirements for the aft skirt thermal conditioning and purge system
are listed in figure 5-70.
The system (figure 5-71) consists of a main manifold which circles
the lox container dome near the aft skirt attach flange, a thruststructure manifold which circles the interior of the thrust cone
structure, a duct connecting the two manifolds, and a hydraulic
accumulator-reservoir manifold attached to the connecting duct.
The system utilizes air as the conditioning medium until 20 min
prior to propellant loading. At this time the conditioning medium
is switched to GN 2 which provides inert purging of the aft skirt
and interstage until liftoff; no conditioning is provided after liftoff.
Compartment conditioning requirements are presented in figure5-72.
The conditioning medium is supplied by the ground ECS and enters
the system through service arm no. 6 umbilical ducting and flows
into the main manifold. A portion of the medium flows into each
of two attitude control modules to maintain the propellant within
5-64
the proper temperature range. Temperature sensors in the gas
discharge of the attitude control modules provide temperature in-
put to the LCC S-IVB ECS panel which controls the conditioning
medium inlet temperature. The remaining portion of the condition-
ing medium flows from the main manifold through orifices to the
aft skirt area for electronic equipment temperature control and
through a duct to the thrust cone manifold and to the hydraulicaccumulator-reservoir shroud. The thrust cone ring manifold dis-
tributes the medium to purge the interior of the thrust structure.
It is then exhausted into the aft interstage area, joining the flow
exhausted from the accumulator-reservoir shroud and flow past
the impingement curtain from the aft skirt area. Slots in the aft
interstage structure then vent the conditioning medium to the
atmosphere. Thermal control of temperature sensitive equipment
during flight is achieved by passive methods. Flight environment
is maintained by controlling preflight temperatures and by control-
ling equipment thermal radiation and conduction paths. Low emis-
sivity, goldized Kapton is used to reduce radiative heat loss and
maintain equipment within allowable limits. A retrorocket impin-
gement curtain between the aft skirt/aft interstage attach flange
and the thrust cone provides protection to electronic components
from hot retrorocket exhaust gases during the separation sequence.
HAZARDOUS GAS DETECTION.
A one-quarter-inch aluminum leak detection manifold mounted
in the forward skirt has four orificed ports for obtaining atmosphere
samples. The manifold connects to a hazardous gas analyzer (HGA)
in the ML through the service arm no. 7 umbilical. A one-quarter-inch aluminum leak detection manifold mounted in the aft inter-
stage has two orificed ports for obtaining atmosphere samples. This
manifold connects to the HGA through the S-IB stage service armIA umbilical. For additional information, see Hazardous Gas De-
tection System in Section IV.
ORDNANCE.
Ordance components perform staging of the Saturn IB vehicle by
severing a tension strap that secures the S-IB and S-IVB stages,by providing the force to decelerate the spent S-IB stage, and by
providing a slight acceleration to the S-IVB stage/payload to keep
the S-IVB stage propellants seated until J-2 engine ignition. Shaped
charges installed on the LH z and lox tanks provide the capability
to destroy the vehicle by severing the tanks and dispersing the
propellants.
STAGE SEPARATION.
S-IB/S-IVB stage separation occurs 1.3 sec after S-IB stage outboard
engine cutoff on command by the S-IB stage switch selector. The
S-IVB stage separates from the S-IB stage/aft interstage at vehicle
station 1186.804 by simultaneous operation of: (1) the stage separa-
tion ordnance system, which severs a circumferential tension plate;
(2) four retromotors, which decelerate the S-IB stage/aft interstageassembly; and (3) three ullage rockets, which maintain a slight
acceleration on the S-IVB stage and payload.
Operation of the stage separation systems begins when the S-IB
stage fuel level sensor no. 1 or no. 2'or lox level sensor no. 2
or no. 3 actuates. Sensor actuation applies 28 Vdc to the chargecircuits of the separation ordnance system and retromotors ignition
system exploding bridgewire (EWB) firing units. At 2.5 sec before
separation, an S-IVB stage switch selector command applies 28
Vdc to the charge circuits of the ullage rocket ignition system EWB
firing units.
Sectian V S-IVB Stage
Figure 5-695-65
Section V S-IVB Stage
REQUIREMENT
SYSTEM LEAKAGE AREA
SYSTEM PROOF PRESSURE
PRELAUNCH PURGE CONTROLS
PRELAUNCH THERMAL CONTROL
AFT SKIRT ELECTRONIC EQUIPMENT ENVIRONMENT
ACCUMULATOR RESERVOIR
ATTITUDE CONTROL MODULE GAS DISCHARGE
INFLIGHT THERMAL CONTROL
Figure 5-70
LIMITS/NOTES
2.5 in. 2 Max.
0.75 pslg
OBTAIN 4% BY
VOLUME 02CONTENT IN
30 MINUTES (Max)
0- 120°/:
0- I00°F
65 -900 F
PASSIVE METHODS
iI
CH.14285.1
At 0.1 sec before stage separation command, the S-IVB stage switch
selector triggers the EBW firing units that fire the ullage rockets
and, at separation, the S-IB stage switch selector triggers the EBW
firing units that fire the retromotors and detonate the separation
ordnance. Approximately 15 sec after stage separation, the spent
ullage rockets and their fairings are jettisoned to reduce stageweight.
Separation Ordnance Subsystem.
The stage separation ordnance system consists of two EBW firingunits, two EBW detonators, a detonator block, and a mild detonat-
ing fuse (MDF) assembly. See figures 5-73 and 5-74.
The MDF assembly, consisting of two parallel lengths of 10-gpf
lead-sheathed PETN enclosed in a polyethylene plastic covering,
is installed in a groove around the aft skirt assembly just forward
of the separation plane. A tension plate, consisting of eight sections
permanently attached to the aft skirt assembly and bolted to the
aft interstage assembly, secures both assemblies until stage separa-
tion command. The MDF assembly lies just'beneath the thinnest
part of the tension plate, which is approximately 0.050 in. thick.
The explosive force created by either MDF will sever the plate;
AFT SKIRT THERMALCONDITIONING
5-66
MAIN
MANIFO_ AIR/GN 2 INLET
APS MODULE_tlm elmemI
_',_ /-,IF/.( \ /__ _'--AFTmi \ 5K'RTlI _ MANIFOLD |
| FEED DI.ICT |
IMPINGEMENT / _ |
_ - I " /';"_t'--THRUST_._I_'K._".\\\ l V/J'JW STR_URE;
LOWERMANIFOLD"--11111111i)III/'ir'r_-___ I
|ll/lll lllll P,o OuT lAFT IIIIIIIIIII II II II II II I; II II iiirllt-_'1',. IINTERSTAGE_ II IIIIli IliW--'_ _'-- HYDRAULIC I
\ IIIl_J II II II II I_ ACCUM-RESERVOmRI
ENGINE |
ITEMPERATURE
INDICATION
TO GSE
,_UMBILICAL SWING ARM NO. 2
CONDITIONING MEDIUM
SUPPLY
CONDITIONED
AIR OR GN 2
Figure 5-71
CH.14335-!
Section V S-IVB Stage
FLOW
MEDIUM
AIR
GN 2
FLOW
(LI_/M;n)
225 - 300
267 - 300
,! I iil:l I:i ill[oJ_',lil';|;I I_'_J|;i IIi[o]'ltl I I lor I r
COMPARTMENT INLET CONDITIONS
PRESSURE
(In. of H20)
l0- 25
10 - 25
TEMP.
(Degrees F)
75- 140
75- 140
HUMIDITY
(Gr,/Lb of Air)
0-43
J 0--1.0
rll :i :[Pill I :i I ",11,'I
COMPARTMENT
TEMP. RANGE
(Degrees F) LOCATION
STAGE
STAGE
Figure 5-72
however, the redundant fuses ensure severance. After installing the
MDF assembly around the aft skirt, each end is secured in acommon detonator block, which contains two EBW detonators. Two
spacers in the detonator block orient the fuse assembly ends with
the detonators. Each detonator fires one end of the MDF assembly,
thereby providing redundant ignition. Electrical cables deliver a
2300 ---+100 Vdc pulse from the EBW firing units to the detonators
at separation command. The detonators fire the MDF assembly,
which propagates the detonation around the stage, severing the
tension plate. Five blast deflectors, mounted just forward of the
separation plane, protect the APS modules and the ullage rockets
from tension plate fragments. The stage separation ordnance igni-
tion components are located in the aft interstage, which remains
with the S-IB stage at separation.
The separation ordnance has successfully met all development and
qualification objectives. Tests included 71 short segments of the
separation joint and three 360-deg, full-scale structural separations,
all without failure. The successful tests of the separation system
and components included: 60 fuse qualification tests, 10 detonator
block tests, six mild detonating fuse installation tests, three full-scale
system tests, and five tests with the system functioning under load.
Retromotor Subsystems.
Four solid-propellant retromotors, mounted at 90-deg intervals
around the aft interstage assembly (figure 5-75), decelerate the S-IB
stage and aft interstage assembly during the stage separation se-
quence. Nose, center, and aft aerodynamic fairings enshroud each
retromotor. The aft and center fairings are permanently installed;
however, the nose fairing jettisons when the retromotor fires to
expose the motor nozzle. Exhaust gases, acting against the internal
surface of the fairing, shear a retaining pin that secures the forward
end of the fairing to the aft interstage assembly. The fairing then
assumes an aftward rotation, pivoting about a hinge and hook
arrangement. After an approximate 70-deg rotation, the fairing
separates from a hook on the center fairing and falls away fromthe vehicle.
Each retromotor nozzle cants 9.5 deg outboard from the motor
centerline to direct the exhaust plume away from the S-IVB stage.
With an average burn time of 1.52 sec, each retromotor developsa 36, 720 lbf thrust "at 200,000 ft altitude and 60 ° F. Thiokol
Chemical Corporation, Elkton, Maryland, manufactures the
E17029-02 recruit motors used as the S-IB stage/aft interstage
assembly retromotors. Each motor weighs 376 Ibm, including 267
Ibm of propellant having a tapered, internal-burning, 5-point-star
configuration.
Independent ignition systems, consisting of two EBW firing units
and two EBW initiators, ignite the retromotors. An igniter, which
is a part of each retromotor, has two threaded receptacles for
initiator installation. A pair of EBW firing units mounts inside
the aft interstage assembly at each retromotor position. An electrical
cable, which is an integral part of each firing unit, connects the
firing unit to its respective initiator. At separation command, eachfiring unit delivers a 2300 -+- 100 Vdc pulse to its initiator. The
initiators detonate and fire the igniter, which directs hot particles
and gases to the solid propellant surface, thereby igniting theretromotor.
N/A
N/A
PROBE DATA
SETTING
(Degree_ F)
65-90
70 - 90
CH.14286-1
To demonstrate satisfactory firing of the motors, tests of the EBW
initiator with the motor igniter were conducted. Eight compatibility
tests were run on the retromotors, including six tests with dual
ignition and two tests with single ignition; all were successful. The
motors were fired under temperature extremes of - 10 ° F to + 155 °
F after being subjected to vibration and temperature cycling tests.
Ullage Rocket Subsystem.
Three solid-propellant ullage rockets, mounted approximately 120
deg apart of the S-IVB stage aft skirt (figure 5-76), induce a slight
forward acceleration to the S-IVB stage and payload during stage
separation. The motors begin firing 0.1 sec before separation and
terminate shortly after J-2 engine ignition. By providing continuous
acceleration to the S-IVB stage during stage separation system
operations, the S-IVB stage propellant remains properly seated inthe bottom of the tanks for J-2 engine start. Aerodynamic fairing
assemblies house the ullage rockets and provide for attachment
to the aft skirt. The fairing assemblies cant each ullage rocket center
line 35 deg outward from the vehicle longitudinal axis, thus direct-
ing the exhaust gases away from the S-IVB stage. Each fairing
assembly also houses a transducer for chamber pressure measure-ment and two EBW firing units for ullage rocket ignition.
SEPARATIONJOINT
"" AFT
SKIRT
O [ _//-"TENSION
]y P'ATE
_''\?.',L_/ SEPARATION _ "_/ III \
Figure 5-73
C-H 14500
5-67
SectionVS-IVBStage
RETROMOTORSANDEBWFIRINGUNITSINSTALLATIONSEPARATION SYSTEM
BLAST DEFLECTOR
ASSY
IV
Ill ASSY
GAP
COVER (8 PLACES)
VIEW LOC)KING AFT
LBLAST
DEFLECTOR
ASSY.
FRAGMENTATION
COVER
GAP COVER
(8 PLACES)
DETONATOR B
EBW FIRING
UNIT NO.
EBW FIRING
UNIT NO. 2
Figure 5-74
Thiokol Chemical Corporation manufactures the TX-280-10 rocketmotors. The motors are 8.316 in. in diameter and a maximum
of 37 in. in length. The internal-burning, 5-point-star-configuration
solid propellant develops an average thrust of 3,460 lbf at a temper-ature of 70 ° F and at an altitude of 1,000,000 ft. At 70 ° F the
propellant burns for 3.9 sec. An igniter, which is part of the rocketmotor and is installed in the forward end of the motor casing,
has two receptacles for EBW initiators. Two initiators, the Thiokol
TX-346-1 and the Aerojet-General AGX 2008, have been approved
for use with the ullage rockets. The output cables from the EBW
firing units in each fairing assembly attach to the respective ullagerocket initiators.
5-68
CENTER
FAIRING-_
AFT
TYPICAL
(4 PLACES) (2 PLACES)
OUTPUT
CABLES
EBW FIRING
UNIT NO.
EBW FIRING /
UNIT NO. 2 OUTBOARD
TYPICAL
(4 PLACES)
C-H 14506
Figure 5-75
The S-IVB stage switch selector issues the ullage rocket ignitioncommand 0.1 sec before stage separation. The command simult-
aneously triggers the two EBW firing units at each ullage rocket
position. Use of redundant EBW firing units and EBW initiators
insures ullage rocket ignition. A 2300 -- 100 Vdc pulse from the
EBW firing units fires the EBW initiators installed in the motor
igniter. The initiators then detonate pellets contained in the igniter.The pellets eject hot particles and gases through perforations in
the igniter case to the solid-propellant surface, igniting the propel-lant.
Thirteen ullage rocket motors were tested in motor qualification,
and all were successful. Eleven were fired utilizing a dual ignition
system, and two were fired utilizing a single ignition system. To
demonstrate EBW initiator/igniter compatibility, six dual ignition
tests and four single ignition tests were conducted successfully. The
testfiringswereconducted under temperature extremes of -30 °
F to + 145 ° F after the motors had been subjected to vibration
tests, temperature cycling tests, and long duration tests of temper-
ature conditioning.
Ullage Rocket Jettison Subsystem.
Approximately 15 sec after stage separation, the ullage rockets and
their fairings are jettisoned to reduce stage weight. To accomplish
this operation, the jettison system uses two EBW firing units, two
EBW detonators, a detonator block, two confined detonating fuse
(CDF) assemblies, six frangible nuts, and three spring-loaded jet-
tison assemblies. See figures 5-77 and 5-78.
The S-IVB stage switch selector issues the ullage rocket jettisoncommand, which triggers the EBW firing units. The 2300 _ 100
Vdc output pulses from the firing units detonate the EBW detona-
tors. The explosion propagates from the detonators through the
CDF assembly, simulatenously breaking all six frangible nuts. The
spring-loaded jettison assemblies then propel the ullage rockets
away from the S-IVB stage.
The EBW firing units, EBW detonators, and detonator block are
mounted at panel position 18 in the aft skirt assembly. A tray
with quick-release clamps, installed on the aft skirt's inner periphery
just forward of the separation plane, secures the CDF assembly
Section V S-IVB Stage
ULLAGEROCKETSJETTISONSYS
EBW FIRING
UNIT NO. 2
EBW FIRING
UNIT NO.
ASSY
ULLAGE ROCKETS AND EBW
FIRING UNITS INSTALLATION
S-IVB
AFT SKIRT
ASSEMBLY_
TX-280- I0
ULLAGE
ROCKET
(3 PLACES
ULLAGE
._51 ULLAGE
ROCKET
59 ° 27' 51'_ __ _
1° 36' 2
ULLAGE
ROCKET
AFT SKIRT ASSY
SKIRT
ING
UNIT NO. 1
UNIT NO. 2
Figure 5-76
ULLAGE
ROCKET
NO. I
/m ULLAGEI ROCKETNO. 3
: -IV
FORWARD _ _j
MOUNTING //1 !', I III I
CKET-- Itll tllCOFASSY-- llllCD CH GE /
FRANGIBLE NUT _._'_
(2 PLACES)_
VIEW A-A
(TYPICAL INSTL)
Figure 5-77
ULLAGE I
ROCKET III
NO. 2VIEW LOOKING AFT AT
VEHICLE STATION 1186.804
CDF ASSY _
ULLAGE ROCKET
AFT MOUNTING
BRACKET
STRINGERS
NO. I1, 48,
&86 "_
VIEW B-B
(TYPICAL INSTL)
C-H 14502
5-69
SectionVS-IVBStage
leads between the detonator block and each ullage rocket position.
A pair of bolts and frangible nuts attach each ullage rocket and
fairing assembly to the aft skirt assembly.
The detonator block contains four threaded receptacles for ordnance
installation: two receptacles, on opposite sides, for the EBW deto-
nators; and two receptacles, on opposite ends, for the CDF as-
semblies. The EBW detonators and CDF assemblies form an X-type
ordnance connection inside the detonator block to provide system
redundancy.. The proximity of the explosives within the detonator
block ensures propagation of the explosion from either (or both)EBW detonator to both CDF assemblies.
Each CDF assembly has three 2-gpf PETN leads with two charge
assemblies on each lead, each charge assembly containing 4.66 gr
of PETN. One end of each lead is adhesively bonded in an end
fitting, which also contains a 2-gr booster charge. The PETN cores
from the three leads butt against the booster charge inside the
end fitting. The end fitting serves two purposes: (1) to install the
CDF assembly in the detonator block and (2) to propagate the
EBW detonator explosion to the PETN core in each lead. Each
CDF assembly lead extends to one ullage rocket position where
the charge assemblies are installed, one in each frangible nut. Two
retaining pins secure the charge assemblies in each nut.
The frangible nuts have a structurally weak plane, which permits
the nut to break open when the charge assemblies detonate. One
charge assembly will fracture a frangible nut; however, the redun-
dant charges ensure nut fracture. After the frangible nuts release
the attachment bolts, a spring-loaded jettison assembly that is
permanently attached to the fairing assembly propels the spent
ullage rocket and fairing assembly away from the stage. Jettisoning
the ullage rockets reduces stage mass by approximately 216 Ibm.
Qualification of the explosive release mechanism for ullage rocket
jettison has been completed. Fifty successful tests of the detonator
block and confined detonating fuse and fittings were completed,
and three complete jettison systems were successfully tested.
Flight History.
The stage separation ordnance flown on the three Saturn IB launch
vehicles performed well within design limits. Flight separation was
completed (S-IVB stage engine clear of the interstage) at approxi-
mately 1.5 sec on AS-201, 1.07 sec on AS-202, and 1.04 sec on
AS-203 after the separation command. The retromotors performed
satisfactorily although AS-201 burn times were longer, and thrusts,
impulses, and pressures were slightly lower than predicted.
Retromotor no. 2 on AS-202 flight experienced a short burn time,
but separation dynamics were well within tolerance. Retromotors
with a larger expansion ratio and a higher thrust rating were flownon the AS-203 flight. The motors were moved 6.75 in. outboard
from the S-IB-I and S-IB-2 retromotor positions. These changes
were made to provide a one-retromotor-out capability for S-IB-3
and subsequent vehicles. Retromotor performance on AS-203 flight
was very close to predicted performance. The ullage rockets per-
formed satisfactorily and jettisoned properly on all three flights.All rocket motors performed within design limits. During the three
flights, J-2 engine thrust increased to the 8000 lbf necessary to
keep the propellants seated before ullage rocket thrust decayedbelow the 8000 lbf minimum.
Firing Units Monitoring.
During checkout operations the EBW tiring units m ttie separation
system, retromotor ignition system, ullage rocket ignition system,
and ullage rocket jettison system are connected to pulse sensorsinstead of the EBW initiators or detonators. A simulated fuel or
lox level sensor actuation command (simulating propellant level
sensor actuation during flight) causes the retromotor and separation
system firing units to charge to 2300 Vdc. Meters on the S-IB EBW
ULLAGE ROCKET
ETTISON ;YSTEM DIAGRAM
5-70
_qD FAIRING ASSY
CDF CHARGE (3 PLACES) ATTACHMENT
ASSY ( 12 PLACES)
UNIT
OUTPUT CABLE
(z PtAC_S)UNIT
(2 PLACES)
C-H 14503
Figure 5-78
ordnance panel provide a readout ot the charge voltage of each
firing unit, which must be 2300 ___ 100 Vdc. A 0 to 4.9 Vdc signal
from each firing unit, proportional to the 0 to 2300 Vdc charge,
operates the meters. The separation command issued by the switch
selector triggers the firing units and the 2300-Vdc pulse discharges
into the pulse sensors. Each pulse sensor provides a continuous
28-Vdc signal to the FIRED indicators (one for each firing unit)on the S-IB EBW ordnance panel. After checkout completion the
pulse sensors are removed from the vehicle and the firing units
output cables are attached to the initiators and detonators. During
flight, the charge voltage of the firing units is telemetered backto KSC and monitored on the S-IB EBW ordnance panel meters.
The measurement numbers for the retromotor firing units are
VM42-400 through VM49-400, and VM68-400 and VM69-400 for
the separation system firing units. These measurements will be
monitored from liftoff until the S-IB stage impacts in the ocean.
Checkout of the ullage rocket ignition and jettison firing units also
requires the use of pulse sensors. The S-IVB stage switch selector
issues the charge and trigger commands to the firing units. Voltage
meters on the S-IVB EBW ordnance panel provide readout of
the charge voltage and FIRED indicators illuminate when the pulsesensors receive the 2300-Vdc pulse from the firing units. During
flight the firing units charge voltage measurements are telemetered
back to KSC and displayed on the S-1VB EBW ordnance panelmeters. Measurement numbers for the ullage rocket ignition firing
units are VM32-416, VM33-416, VM34-417, VM35-417, VM36-418,
and VM37-418. For the ullage rocket jettison firing units, themeasurement numbers are VM38-404 and VM39-404. These mea-
surements are monitored from liftoff through S-IVB burn.
PROPELLANT DISPERSION SYSTEM.
The S-IVB stage propellant dispersion system (PDS) severs the
cylindrical portion LH z tank between stations 1276 and 1536, andthe bottom of the lox tank at station 1144 if flight termination
becomes necessary (figure 5-79). The propellants disperse away
Section V S-IVB Stage
S&A
DEV
UNITS
_ PRIMACORD
FORWARD SKIRT
FUSE ASSY
\LH 2 TANK
LINEAR
SHAPED
CHARGE
PRIMACORD
AFT DOME
FUSE ASSY
LOX TANK
LINEAR
SHAPED
CHARGE
ADHESIVE
FIRING UNIT
OUTPUT CABLES
DETONATORS
EBW
S&A DEVICE "
INSERTS _- PRIMACORD
END FITTINGS
|0--GRAIN PETN
LSC-TO-LSC END FITTINGS (TYPICAL)
ADHESIVE
BOOSTER CHARGE
10--GRAINS PETN
Figure 5-79
BOOSTER CHARGE
6-GRAINS PETN
PRIMACORD-TO-LSC END FITTINGS (TYPICAL)
C-H 14474
5-71
SectionVS-IVBStage
from each other, thus reducing the yield of the burning propellants.
Two EBW firing units, two EBW detonators, a safety and arming
device, a forward skirt fuse assembly, aft dome fuse assembly, LH
tank linear shaped charge (LSC) assemblies, and lox tank, LSC
assemblies make up the redundant S-IVB stage PDS. To fire thePDS ordnance, the EBW firing units deliver a high-voltage high-
energy pulse to the EBW detonators. The detonators then propagate
the explosion through the two leads in the S&.A device and sub-
sequently to the LSC.
Tests that demonstrate the ability of the 150-gpf LSC to cut the
LH 2 tank and lox tank at cyrogenic temperatures have been suc-
cessfully completed. Eighteen shaped-charge tank-cutting tests,
eighteen end-fitting propagation tests, and one safe-arm deviceinstallation test were all successful. The continued reliability of
the ordnance is assured by a 25 percent lot sampling of the compo-
nents. A single failure would cause rejection of the lot.
EBW Firing Unit.
See section iv, S-IB Stage Propellant Dispersion System.
EBW Detonator.
See Section IV, S-IB State Propellant Dispersion System.
Safety and Arming Device
See Section IV, S-IB Stage Propellant Dispersion System.
Forward Skirt Fuse Assembly.
The forward skirt fuse assembly propaga_tes the detonation from
the S&A device to the LH z tank LSC. Two 60-gpf Primacord leads,
each approximately 155 in. long, form a redundant explosive train.
Each lead has a fitting on each end to permit connection of the
RDX CHARACTERISTICS
DESCRIPTION
CRYSTAL DENSITY
CHEMICAL FORMULA
MOLECULAR WEIGHT
MELTING POINT
DETONATION RATE
HEAT OF COMBUSTION
FiRE HAZARD
ICC CLASSIFICATION
SYNONYMS
SPECIFICATION
IMPACT SENSITIVITY
WHITE CRYSTALLINE POWDER
1.816
C3H6N60 6
222.15
204eC
8,180 M/SEC AT 1.65 GM/CC
(I-INCH DIA SAMPLE)
2307 CALORIES/GM
MODERATE, BY SPONTANEOUS
CHEMICAL REACTION
CLASS A
CYCLONITE, HEXAGON
MIL-R-00398
GREATER THAN 120 FOOT-POUNDS.
100 GPF LSC TEST (DAC REPORT
SM-42678), NO DETONATIONOCCURRED WHEN IMPACTED WITH
A 10-L8 WEIGHT FROM A 12-FOOT
HEIGHT.
Figure .5-80
C-H 14477
fuse assembly to the S&A device and LH z tank LSC. A 6-gr PETN
booster charge contained in each end fitting assures propagationacross the mechanical connections. Each of the two end fittings
connecting the fuse assembly to the S&A device consists of a sleeve,in which the Primacord lead and booster charge are adhesively
bonded, and a knurled nut that secures the lead in the S&A device.
The two end fittings that interface with the LH 2 tank LSC arethreaded sleeves. The Primacord leads are wrapped together in
aluminum tape along their entire length. Quick-release clamps
secure the fuse assembly along the power distribution mounting
assembly and mounting panel no. 31 inside the forward skirt, and
along the forward-skirt tunnel-equipment panel assembly installedon the forward skirt exterior under the tunnel.
LH= Tank LSC.
The LH 2 tank LSC installation consists of eight sections of 150-gpf
cyclotrimethylene trinitramine (RDX) enclosed in an aluminumsheath. Each RDX section has bonded end fittings that contain
10-gr booster charges. Redundant explosive trains consisting of fourRDX sections connected by the end fittings bffect LH z dispersion
by severing the tank skin. The LSC design confines the explosive
force to a very narrow path along the tank skin to produce the
cutting effect. The use of different LSC lengths staggers the end-til-ting connections along the LH 2 tank to assure that the proximity
of the parallel LSC assemblies will produce a continuous cut be-tween stations 1276 and 1536. If the end fittings were installed
side by side, the LSC could fail to completely sever the tank skin
underneath the end fittings. Quick-release clamp assemblies, which
are adhesively bonded to the LH 2 tank (under the tunnel), securethe LSC assemblies against the skin. In addition to severing the
LH z tank skin, the LSC propagates the detonation to the aft dome
fuse assembly through end fittings at station 1276. See figure 5-80for RDX characteristics.
Aft Dome Fuse Assembly.
Two 60-gpf Primacord leads, each approximately 250 in. long, form
the redundant explosive train between the LH 2 tank LSC and
the lox tank ISC. Threaded end fittings containing 6-gr PETN
booster charges are adhesively bonded to the ends of the Primacord,
providing the means of attachment to the LH 2 and lox tank LSC
end fittings. Quick-release clamps install the fuse assembly on a
panel underneath the tunnel on the aft skirt assembly and alongthe lox tank aft dome and thrust structure. Under the tunnel, the
fuse assembly is wrapped with aluminum tape. Along the aft dome
and thrust structure, the fuse assembly has three counterwrapped
layers of asbestos tape with an outer layer of glass fiber tape for
protection against heat from the J-2 engine and retromotor exhaust
plumes. The aft dome fuse assembly interfaces with the lox tankLSC at vehicle station 1144 (approx). See Section IV, Ordnance,for PETN characteristics.
Lox Tank LSC.
The lox tank LSC installation consists of three 150-gpf RDX sec-
tions joined together by end fittings. The LSC encircles the base
of the lox tank on a 23-in. radius at vehicle station 1144 (approx).
The end fittings each contain a 10-gr booster charge to ensure
propagation of the detonation across the mechanical connections.Quick-release clamp assemblies that are adhesively bonded to the
lox tank aft dome secure the LSC.
5-72
SECTION VI
TABLE OF CONTENTS
Introduction ...................................................................... 6-1
Structure ............................................................................ 6-1
Environmental Control System ........................................ 6-2
Electrical Power Systems ................................................ 6-10
Emergency DetecUon System .......................................... 6-12
Launch Vehicle Navigation,
Guidance, and Control ................................................ 6-13
Measurement and Telemetry .......................................... 6-24
Command System ............................................................ 6-25Saturn Tracking Instrumentation .................................... 6-25
Ground Support Equipment ............................................ 6-27IU/SLA Interface .............................................................. 6-28
Experiments ...................................................................... 6-29
INTRODUCTION.
The Instrument Unit (IU) is a cylindrical structure installed on
top of the S-IVB stage (figure 6-1). The IU contains the guidance,
navigation, and control equipment which will guide the vehicle
through its launch trajectory and subsequently to provide attitude
control during orbital operations. In addition, it contains telemetry,
communications, tracking, and crew safety systems, along with their
supporting electrical power and environmental control systems.
This section of the Flight Manual contains a description of the
physical characteristics and functional operation for the equipmentinstalled in the IU.
STRUCTURE.
The basic IU structure is a short cylinder fabricated of an aluminum
alloy honeycomb sandwich material (see figure 6-2). The top and
bottom edges are made from extruded aluminum channels bonded
to the honeycomb sandwich. This type of construction was selected
for its high strength-to-weight ratio, acoustical insulation, and ther-
mal conductivity properties. The cylinder is manufactured in three
120 degree segments which are joined by splice plates into an
integral structure. The three segments are the access door segment,
the flight control computer segment, and the ST-124M-3 segment.
The access door segment has an umbilical door, as well as al,
equipment/personnel access door. The access door has the require-ment to carry flight loads and still be removable at any time prior _1
to flight. luAttached to the inner surface of the cylinder are cold plates which
serve both as mounting structures and thermal conditioning units
for the electrical/electronic equipment. Mounting the electrical/
electronic equipment around the inner circumference of the IU
.Ill
VHF
TELEMETRY
ANTENNA
(2 PLACES)
ANTENNA
(2 pLACES)
COMMAND
ANTENNA
(2 PLACES)
Figure 6-1
UMBILICAL
DOOR
CH.14479.2
6-1
Section Vl Instrument Unit
LAMINATED HONEYCOMB SANDWICH MATERIAL
C-H 20_@
Figure 6-2
leaves the center of the unit open to accommodate the convex
upper tank bulkhead of the S-IVB stage.
Cross section "A" of figure 6-3 shows equipment mounting pads
bolted and bonded to the honeycomb structure. This method is
used when equipment is not mounted on thermal conditioning cold
plates. The bolts are inserted through the honeycomb core, and
the bolt ends and nuts protrude through the outside surface. Cross
section "B" shows a thermal conditioning cold plate mounting panelbolted to brackets which, in turn, are bolted to the honeycomb
structure. The bolts extend through the honeycomb core with the
bolt heads protruding through the outer surface. Cross section "C"
shows the cable tray supports bolted to inserts, which are potted
in the honeycomb core at the upper and lower edges of the structure.
Figure 6-4 shows the relative locations of all equipment installedin the IU.
ENVIRONMENTAL CONTROLSYSTEM.
The Environmental Control System (ECS) maintains an a_ceptable
operating environment for the IU and S-IVB forward skirt equip-
ment during preflight and flight operations. The ECS is composed
of the following:
a. The thermal conditioning system (TCS) which maintains a cir-
culating coolant at a controlled temperature to the cold plates and
certain internally-cooled components mounted on the IU and theS-IVB forward skirt.
b. The preflight purging system which maintains a supply of tem-
perature and pressure related air/GN_ to the IU/S-IVB equipmentarea.
c. The gas-bearing supply which furnishes GN 2 to the ST-124M-3
inertial platform gas bearings.
THERMAL CONDITIONING SYSTEM (TCS).
Up to sixteen thermal conditioning panels; (cold plates), each
capable of absorbing up to 420 watts of thermal energy, may be
located in each of the IU and S-IVB stages. Each cold plate contains
tapped bolt inserts in a fixed grid pattern which provide flexibility
of component mounting. Temperature control is accomplished by
6-2
circulation of a coolant fluid, Oronite Flocool 100", through pas-
sages in the cold plates.
A functional flow diagram is shown in figure 6-5. Two heat ex-
changers are employed in the system. One is used during the
preflight mode and employs GSE supplied circulating Oronite as
the heat exchanging medium. The other is the flight mode unit,which uses demineralized water and the principle of sublimation
to effect heat dissipation.
The manifold, plumbing, and both accumulators are filled during
the prelaunch preparations. The accumulators serve as positive
pressure reservoirs supplying fluid to their respective systems ona demand basis. There is a flexible diaphragm in each accumulator,
backed by regulated low pressure GN 2.
During operation of the TCS, the Oronite coolant is circulated
through a closed loop by an electrically driven centrifugal pump
(a secondary pump provides redundancy). The supply manifold
diverts part of the coolant to the cold plates in the S-IVB stage
and the remainder to the cold plates, gas bearing heat exchanger,
inertial platform, LVDC/LVDA, and flight control computer inthe IU.
During the preflight mode, the sublimator, which functions onlyin a vacuum environment, is inactive and a solenoid valve blocks
the water flow. The preflight heat exchanger transfers heat from
the closed loop fluid to GSE fluid.
Approximately 180 seconds after liftoff, the water solenoid valve
is opened and the sublimator becomes active. During the periodbetween GSE disconnect (at liftof0 and sublimator activation, the
capacity of the pre-cooled system is sufficient to preclude equipment
overheating.
The sublimator element is a porous plate. Since the sublimator
is not activated until approximately 180 seconds after launch, the
ambient temperature and pressure outside the porous plates are
rapidly approaching the vacuum conditions of space. Water flows
readily into the porous plates and into the pores. The water freezes
when it meets the low temperature and pressure of the space
environment, and the resulting ice formation blocks the pores
(figure 6-6).
As heat is generated by the equipment, the temperature in theOronite solution rises. This heat is transferred within the sublimator
to the demineralized water. This heat is then dissipated to space
through the latent heat of sublimation as the ice is transformed
to vapor by the space environment. The vapor is vented into the
IU compartment. As the heat input decreases, the rate of sublima-
tion is reduced, decreasing the water flow. Thus, the sublimator
is a self-regulating system. However, if the coolant temperature
falls below the lower limit of approximately 59 ° F, a switch selectorcommand is issued to close the water solenoid valve causing the
cessation of sublimator operation. GN 2 for the Oronite and water
accumulators is stored in a 165 cubic inch sphere in the IU at
a pressure of 3,000 psig. The sphere is filled prior to liftoff by
applying high pressure GN 2 through the umbilical. A solenoid valve
controls the flow into the sphere, and a pressure transducer indicates
to the GSE when the sphere is pressurized. The GN 2 flow from
the sphere is filtered and applied to the accumulators through a
pressure regulator, which reduces the 3,000 psig to 16.5 psia. An
orifice regulator further reduces the pressure at the water accumu-
lator to approximately 5 psia. The GN 2 flow is then vented within
the IU compartment.
PREFLIGHT AIR/GN= PURGE SYSTEM.
The preflight air/GN 2 purge system directs ground supplied, tem-
perature and pressure regulated, filtered air or GN_ to the IU/S-IVB
*Reg. TM, Chevron Chemical Co,
Section VI Instrument Unit
INSTRUMENTUNIT STRUCTURALDETAILS
CA,LETRAY
SUPPORTS
HORIZONTAL
SPLICE PLATE -_
INNER
SPLICE PLATE _ f
OUTER
SPLICE PLATE _
_J
TYPICAL SPLICE PLATE
ATTACHMENT
0.200
.IN.
SLA INTERFACE
UPPERMOUNTING
RING
_-_ i. 0,95 IN.
'_ FULL WALL
THICKNESS
KX×_XK_<X _ LOW DENSITY
_ CORE
_ _0030ALUM7075-T6
I_I" OUTERSKINIR;:_:_I / ALLOY
J_ 7075-T6INNER SKIN
II;5_<5;_I _ LOWERI1_:_1 / _OUNTING
_l_ RING
0.125
'L-----t- _----.4
1.25 IN.
S-IVB I NT ERFACE
36 IN.
TYPICAL SECTIONAL VIEW
Figure 6-3
C.H 20010
6-3
Section VI Instrument Unit
IU COMPONENTS
POWER
DISTRIBUTOR
ELECTRONIC
CONTROL ASSY601A40
(PITCH) 601A25
POWER
601A401
56 VOLT POWER
SUPPLY 601A15
MEASURING RACK
SELECTOR
601A400
SA-206, 207 & 209
RACK
FLIGHT
IV-I
i /
_ IV
II_iii
LOCATION VIEW
POWER
DISTRIBUTOR
ENVIRONMENTAL
CONDITIONING
DUCT
POWER
DISTRIBUTOR
601A33
601A401
RACK
SENSOR 603AB2
POWER
ELECTRONIC SUPPLY 601A15
CONTROL ASSY
601A40 MEASURI NG RACKSELECTOR
601A400
CONTROL ACCELEROMETER
(PITCH) 60 IA25
SA-208 ONLY
Figure 6-4 (Sheet 1 of 6)
C-BAND
RADAR
TRANSPONDER
601A635
LOCATION VIEW
C.H 20011
Section Vl Instrument Unit
U COMPONENTS
BATTERY ASSY 601A9
BATTERYI'L_ ILS_/_,_ _ _ WATER
COLD PLATE _ t
ACCUMULATOR
_ _Y ASSY _ TELEMETER POWER
g,fe.'_ _ 601A10 DIVIDER 601A436
_ RF DUMMY%. LOAD 601A569 I
_-_0*#Lsw,;_H /
II ___ IV
FLIGHT LOCATION VIEW
EXPANSION CHAMBER
ENVIRONMENTAL
GAS BEARING
SOLENOID VALVE
)LANT PUMP
(PRI MARY)601A37
CHECK
VALVE (2 PLACES)
PUMP
(SECONDARY)601A62
ACCESS DOOR
UMBILICAL
DOOR
S-IVB
RETURN
GN 2 SPHEREASSY (165 IN 3
HEAT
EXCHANGER ASSY
Figure 6-4 (Sheet 2 of 6)
VALVE 601A41
SUPPLY
III _"_"_ I
LOCATION VIEW
C-H 14471.2
6-5
Section VI Instrument Unit
IU COMPONENTS
MEASURING
DISTRIBUTOR
602A4
(SA-206 ONLY).
MEASURI NG RACK
602A405
ENVIRONMENTAL CONDITIONING DUCT
EDS DISTRIBUTOR
602A420
602A634
ADAR TRANSPONDER
tFLIGHT
MEASURING
RACK 602A406
(SA-206 ONLY)
SA-206, 207, & 209
TELEMETER F1
602A439
METER RF ASSY
FI602A643
MASTER MEASURING VOLTAGE SUPPLY
602A16
IV _I._ I II
LOCATION VIEW
P10 GAS
STORAGE
SPHERE
602A420
C-BAND RADAR TRANSPONDER
602A634
FILTER -
GAS BEARING
PRESSURE
REGULATOR
'/ENT VALVE
RLL& VENT VALVE
SHUTOFF VALVE
TELEMETER RF ASSY P1
602A600
SA-208 ONLY
rELEMETER FI
602A439
EMETER RF ASSY
F1602A643
602A16
PPLY
C-N 20012
6-6 Figure 6-4 (Sheet 3 of 6)
SectionVIInstrumentUnit
IU COMPONENTS
RONMENTAL CONDITIONING DUCT
RATE GYRO
PACKAGE
602A23
SWITCH SELECTOR
603A17
4OSEC iTIMER
603AS6_
COMMAND POWEF
DIVIDER 603A451
tFLIGHT
DDAS COMPUTER
INTERFACE UNIT
603A447
COMMAND DIRECTIONAL
COUPLER 603A562
COMPUTER 602A27
SA-206, 207, & 209
3NAL
PROCESSOR 602A24
RACK
602A4O4
MEASURING RACK
602A4O3
III
/
IV___I I
I
LOCATION VIEW
DDAS COMPUTER
INTERFACE UNIT
RATE GYRO
PACKAGE
602A23
SWITCH SELECTOR
603A17
_-NAL
PROCESSOR 602A24
4O SEC
TIMER
602A4O4
COMMAND POWER'
COMMAND DIRECTIONAL
COUPLER 603A562
FLIGHT CONTROL
COMPUTER 602A27
EXPERIME t_rT CONTROL
DISTRIBUTOR (S- 150)
603A88
SA-208 ONLY
Figure 6-4 (Sheet 4 of 6)
MEASURING RACK
602A4O3
III
/
I
LOCATION VIEW
C-H 20013
6-7
Section VI Instrument Unit
IU COMPONENTS
DIGITAL
MULTI PLEXER
MOD 410
602A615
MULTIPLEXER
/ ASSY P1
DISTRIBUTOR / 602A446602A3
RACK 602A402
DUCT)NMENTAL CONDITIONING
IXILIARY POWER DISTRIBUTOR
602A34
_LEXER ASSY F2
MOO 270 602A575
MOO II
TM CALIBRATOR
602A644
TELEMETER RFASSY P1 602A_71
(SA-206, 207, & 209 ONLY)
TM CALIBRATOR
POWER & CONTROL
ASSY 602A566
TM DIRECTIONAL
COUPLER 602A668
tFLIGHT
II
/
III__ _ I
IV
LOCATION VIEW
ST-124M-3
PLATFORM
ELECTRONIC ASSY
4MENTAL
CONDITIONING DUCT
PLATFORM
AC POWER
SUPPLY 603A13
ACCELEROMETER
SIGNAL CONDITIONER
503A21
_ LVDA 603A29
CONTROL DISTRIBUTORF 603A2
,/,
COMMAND DECODERREMOTE DIGITAL 603A450MULTIPLEXER
MOD 410
603A455LVDC 603A28
Iv_
I _/Ill
LOCATION VIEW
C-H 2001#,
Figure 6-4 (Sheet 5 of 6)6-8
Section Vl Instrument Unit
IU COMPONENTS
GAS BEARI NG
SOLENOID VALVE _ CONTROL
603A43 _ ACC ELFROM ET ER
/ 603A26
/'--" ENVIRONMENTAL / _-------'_ _-
r CONDITIONING DUCT++_ _ _r'"----"-_l
"_"_]" --I _ _ _'-_ /.--- LSNGITUDINAL ACC ELEROMETER
_"_ _'_ n _ _k_ __ GAI4BEoARING HEAT EXCHANGER
_iBADA_s:OND ER / It _ ___ _
603A635 _ i __S_BBT::,I_: FII:EERLEROMET ER
y_ \ _,_ _j _ _ _ ST-124M-3 STABILIZED PLATFORM ASSY60_Al?
_ GAS BEARING REGULATOR
_ SPHERE ASSY IV-I IV
\ /
SA-206, 207 & 209 _ III
tFLIGHT
ASAP MEMORY
ASSY
ASAP DC-DC
CONVERTER
ASAP TAPE
RECORDER
CONTROLACCELEROMET ER
ENVIRONMENTAL
_NGITUDINAL ACCELEROMETER
_ _ /_-_. GAS BEARING HEAT EXCHANGER
41 _\\tl _ A _ _ /_/p_/_"_JJ _ LONGITUDINAL ACCELEROMETER
GAS BEARING FILTERIII 7/--.._ _'_'__ _:_I_ /a
• IV
_ GAS BEARING REGULATOR IV-I /
_COMPUTERINTERFACE SPHERE ASSY I
UNiT 603A84 __ _ lil
II LOCATION VIEW
UNIT 603A83
SA-20BONLY C-. 20015
Figure 6-4 (Sheet 6 of 6) 6-9
SectionVl Instrument Unit
interstage compartment. The air or GN 2 is distributed through a
flexible duct system mounted above the payload interface as shown
in figure 6-1. Ventilating air for temperature and humidity control
is furnished during preflight phases. During fueling, inert GN sis furnished to prevent the accumulation of a hazardous and corro-
sive atmosphere. The air or GN 2 flows through holes in the ductingpreviously described.
GAS BEARING SUPPLY.
Gaseous nitrogen, for the ST-124M-3 inertial platform, is stored
in a two cubic foot sphere in the IU at the pressure of 3,000 psig
(figure 6-5). The sphere is filled by applying high pressure GN zthrough the umbilical under control of the IU pneumatic console.
A low pressure switch monitors the sphere; and, if the pressurefalls below 1,000 psig, the ST-124M-3 stable platform is shut down
to preclude damage to the gas bearings during ground checkoutoperations. This switch is inactive during flight.
Output of the sphere is through a filter and a pressure regulator.The regulator reduces the sphere pressure to a level suitable for
gas bearing lubrication. Pressure internal to the platform is sensed
and applied as a control pressure to the regulator. This provides
for a constant pressure differential across the gas bearings. The
gas flows from the regulator through a heat exchanger, where its
temperature is stabilized, then through another filter and on to
the gas bearings. Spent gas is then vented into the IU compartment.
ELECTRICAL POWER SYSTEMS.
Primary flight power for the IU equipment is supplied by three
(3) silver-zinc batteries at a nominal voltage level of 28(--+2) Vdc.
During prelaunch operations, primary power is supplied by the
GSE. Where ac power is required within the IU, it is developed
by solid state dc-to-ac inverters. Power distribution within the IU
is accomplished through power distributors which are, essentially,
junction boxes and switching circuits.
BATTERIES.
Silver-zinc primary flight .batteries are installed during prelaunch
operations at the locations shown in figure 6-4, sheet 2. These
batteries are identical, each having the characteristics shown in
figure 6-7. Each battery is connected to a separate bus in a powerdistributor. The D I0 and D30 batteries are connected to a common
bus through isolation diodes to provide a redundant power source
for critical IU platform, switch selector, and control functions. Flight
components are connected to the buses in the various distributors.
The silver-zinc batteries are characterized by their high efficiency.
Their ampere-hour rating is about four times as great as that of
a lead-acid or nickel-cadmium battery of the same weight. Thelow temperature performance of the silver-zinc batteries is also
substantially better than the others.
THERMAL CONDITIONING SYSTEM DIAGRAM
PREFLIGHT AIR/GN 2 PURGE SUBSYSTEM
PRESSURE
SWITCH
SNUBBER
ASSEMBLY
COOLANT TEMPERATURE
ORIRCE ASSEMBLY
GAS BEARING HEAT ,ST-124M INERTIAL ,LAUNCH VEHICLE
REDUNDANT
GSE
AIR OR GN 2THERMAL
EXPANSION
CHAMBERUMBILICAL
GSE I
SUPPLIED
COOLANT (O
VAPOR
VENT
ACCUMULATOR
RETURN _ GN 2SOLENOID
COLD VALVE BPLATE
(16 REQ'D)
GAS BEARING LOW PRESSURE
. PREFLIGHT PRESSURE SWITCHHEAT
EXCHANGER REGULATOR
WATER RRST STAGE (2 FT3) GN2SOLENOID
R LL PRESSURE GN 2 STORAGE SPHERE VALVE A
REGULATOR
FLIGHT HEAl WATER
EXCHANGER : ACCUMULATOR
(SUBLIMATOR) = t
GN 2 SUPPLY &_.-._.EMERGENCY _ "_VENT i
SHUT-O FF VALVE
LEGEND
COOLANT
r/'/-r/TI COOLANt (GSE SUPPLIED)
ll_ GASEOUS NITROGEN (GN2)
WATER AND WATER VAPOR
CALIBRATE LINE
ORIRCE
REGULATOR
FILTER A
MEMBRANE165 IN.3
BLEEDER GN 2 STORAGE SPHERE
ASSEMBLY
DISCONNECT (TYPICAL)
PLATE
(5 REQ'D)
INSTRUMENT UNIT
S-IVB STAGE
RETURN
6-10
Figure 6-5
CONTROL
COMPUTER
HICLE
DATA ADAPTER
RETURN FLOW
C.H 29120
Section VI Instrument Unit
SUBLIMATORDETAILS
VACUL NT PASSAGE
SURFACE
TENSION
COOLANT
WATER PASSAGE
POWER CONVERTERS.
The IU electrical power systems contain a 56-V power supply and
a 5-V measuring voltage supply.
56-V Power Supply.
The 56-V power supply furnishes the power required by the ST-
124M-3 platform electronic assembly and the accelerometer signal
conditioner. It is basically a de-to-de converter that uses a magnetic
amplifier as a control unit. It converts the' unregulated 28 Vdc
from the batteries to a regulated 56 Vdc. The 56-V power supply
is connected to the platform electronic assembly through the power
and control distributors.
IU BATTERY CHARACTERISTICS5-V Measuring Voltage Supply,
The 5-V measuring voltage supply converts unregulated 28 Vdc
to a closely regulated 5 (+.005) Vdc for use throughout the IU
measuring system. This regulated voltage is used primarily asexcitation for measurement sensors (transducers), and as a reference
voltage for inflight calibration of certain telemetry channels. Like
the 56-V supply, it is basically a de-to-de converter.
DISTRIBUTORS.
The distribution system within the l U is comprised of the following:
1 Measuring distributor (2 used on SA-206 only)
I Control distributor
1 Emergency Detection System (EDS) distributorI Power distributor
2 Auxiliary power distributors
1 Experiment distributor (used on SA-208 only)
POROUS PLATE
C-H 20016
Figure 6-6
Measuring Distributors.
The primary function of the measuring distributors is to collectall measurements that are transmitted by the IU telemetry system,
and to direct them to their proper telemetry channels. These mea-surements are obtained from instrumentation transducers, func-
tional components, and various signal and control lines. The mea-
suring distributors also distribute the output of the 5-V measuring
voltage supply throughout the measuring system.
Through switching capabilities, the measuring distributors can
change the selection of measurements monitored by the telemetry
TYPE
MATERIAL
CELLS
NOMI NAL VOLTAGE
ELECTROLYTE
OUTPUT VOLTAGE
OUTPUT CURRENT
GROSS WEIGHT
DRY CHARGE
ALKALINE SILVER-ZINC
20 (WITH TAPS FOR SELECTING
18OR 19 CELLS IF REQUIRED
TO REDUCE HIGH VOLTAGE)
1.5 PER CELL
POTASSIUM HYDROXIDE (KOH) IN
DEMINERALIZE D WATER
+28 ± 2 VDC
35 AMPERES FOR A 10 HOUR LOAD
PERIOD (IF USED WITHIN 120
HOURS OF ACTIVATION)
165 FOUNDS EACH
Figure 6-7
C.H 2_17
6-11
Section Vl Instrument Unit
system. The switching function transfers certain measurements to
channels which had been allotted to expended functions. If it werenot for this switching, these channels would be wasted for the
remainder of the flight.
Control Distributor.
The control distributor provides distribution of 28-V power to small
current loads and distributes 56 Vdc from the 56-V power supplyto the ST-124M-3 inertial platform assembly. The control distrib-
utor provides power and signal switching during prelaunch checkout
for testing various guidance, control, and EDS functions, requestedby the launch vehicle data adapter through the switch selector.
Emergency Detection System Distributor.
The EDS distributor provides the only electrical link between the
spacecraft and the LV. All EDS signals from the LV are routed
to the logic circuits in the EDS distributor. EDS output signalsfrom these logic circuits are then fed to the spacecraft and to the
IU telemetry. Also, EDS signals from the spacecraft are routed
back through the IU EDS logic circuits before being sent to theS-IVB and S-IB stages.
Power Distributor.
The power distributor provides primary distribution for all 2_-V
power required by IU components. Inflight 28-V battery power,
or prelaunch ESE-supplied 28-V power, is distributed by the powerdistributor as shown in figure 6-8.
The power distributor also provides paths for command and mea-
surement signals between the ESE and IU components. The power
distributor connects the IU component power return and signal
return lines to the IU single point ground and to the umbilicalsupply return bus. These return lines are connected to the common
bus in the power distributor, directly or indirectly, through oneof the other distributors.
Auxiliary Power Distributors.
Two auxiliary power distributors supply 28-Vdc power to small
current loads. Both auxiliary power distributors receive 28 Vdc from
each of the battery buses in the power distributor. Relays in the
auxiliary power distributors provide power ON/OFF control for
IU. components during the prelaunch checkout. These relays arecontrolled by the ESE.
IU GROUNDING.
All IU grounding is referenced to the outer skin of the LV. The
power system is grounded by means of wires routed from the powerdistributor COM bus to a grounding stud attached to the LV skin.
All COM buses in the various other distributors are wired back
to the COM bus in the power distributor. This provides for a singlepoint ground.
Equipment boxes are grounded by direct metal-to-metal contact
with cold plates or other mounting surfaces which are common
to the LV skin. Most cabling shields are grounded to a COM bus
in one of the distributors or to the equipment case.
During prelaunch operations, the IU and CSE COM buses are
referenced to earth ground. To ensure the earth ground reference
until after all umbilicals are ejected, two single-wire grounding
cables are connected to the IU below the umbilical plates. Theseare the final conductors to be disconnected from the IU.
EMERGENCY DETECTION SYSTEM.
The EDS is the principle element of several crew safety systems.
TYPICAL IU POWER DISTRIBUTION
6-12
BATTERY J6D10
1I' I ' ' I '6Dll 6D41 6Dll 6D41
DISTRIBUTOR DISTRIBUTOR
601A33 602A34
I I
PRIMARy" BUS F_OWERTO COMPONENTS
REQUIRING UP TO
10 AMPS
5 vo_
MEASURING
VOYAGE
SUPPLY
6D81 J
MEASURING
DISTRIBUTOR
602A4
I BATTERY6D30
LPOWER DISTRIBUTOR601AI
6Dll J 6D41
6D31
1CONTROL
DISTRIBUTOR
603A2
J BUS POWERTO LOW
CURRENT
COMPONENTS
I BATTERY I6D40
t6D40
1
J I COOLANT
PUMP NO. I
AND NO, 2
l
6D51
CONTROL FUNCTIONS
J,o l o'i'°,L'i2
I
NOTE: ONLY POSITIVE BUSES
ARE SHOWN
C-H 200]8
Figure 6-8
EDS design is a coordinated effort of crew safety personnel fromseveral NASA centers.
The EDS senses development of conditions which could ultimately
cause vehicle failure. The EDS reacts to these emergency situations
in either of two ways. If breakup of the vehicle is imminent, an
automatic abort sequence is initiated. If, however, the emergency
condition is developing slowly enough, or is of such a nature that
the flight crew can evaluate it and take action, only visual indica-
tions are provided to the flight crew. Once an abort sequence hasbeen initiated, either automatically or manually, it is irrevocable
and runs to completion.
The EDS is comprised of sensing elements, signal processing and
switching circuitry, relay and diode logic circuitry, electronic timers
and display equipment, all located in various places on the flight
vehicle. Only that part of the EDS equipment located in the IU
will be discussed here.
There are nine EDS rate gyros installed in the IU. Three gyrosmonitor each of the three axes (pitch, roll, and yaw) thus providing
triple redundancy.
The control signal processor provides power to the nine EDS rate
gyros, as well as receiving inputs from them. These inputs are
processed and sent to the EDS distributor and to the flight control
computer.
The EDS distributor serves as a junction box and switching device
to furnish the spacecraft display panels with emergency signals
if emergency conditions exist. It also contains relay and diode logic
for the automatic abort sequence.
There is an electronic timer, which is activated at liftoff and which
produces an output 40 seconds later. This output energizes relaysin the EDS distributor which allows multiple engine shutdown,
which had been inhibited during the first 40 seconds of launch.
Inhibiting of automatic abort circuitry is also provided by the LVflight sequencing Circuits through theIU switch selector. This inhib-
iting is required prior to normal S-IB engine cutoff and other"normal LV sequencing. While the automatic abort capability is
inhibited, the flight crew must initiate a manual abort, if an angu-
lar-overrate or two-engine-out condition occurs.
See Section III for a more complete discussion of emergency detec-
tion and procedures. Section 1II includes launch vehicle monitoringand control, EDS controls, and abort modes and limits.
LAUNCH VEHICLE NAVIGATION,GUIDANCE, AND CONTROL.
The Saturn 1B astrioni'cs system will provide navigation, guidance
and control (N, G, & C) of the launch vehicle from liftoff until
separation of the Apollo spacecraft from the S-IVB/IU and attitudecontrol until the end of active lifetime. The function of the N,
G, & C portion of the astrionics system (figure 6-9) is to steerthe launch vehicle along an optimum trajectory (minimum fuel
consumption and safe structural loading) into a pre-determinedearth orbit. Orbital N, G, & C functions include computations to
solve for position and velocity (based on insertion conditions) andvehicle attitude control. Figure 6-10 defines the N, G, & C functions
and the equipment or program that will perform the functions.Other functions performed by or interfacing with the N, G, &
C hardware are event sequencing and telemetry-data management.Alternate altitude error commands can be provided to the control
system from the Apollo spacecraft in the event of a guidancereference failure during any flight phase. During the orbital phase,
the spacecraft may provide control system input regardless of the
guidance system health.
Section Vl Instrument Unit
N, G, AND C SUMMARY.
The ST-124M inertial platform system establishes a space-fixed
reference coordinate system and provides instantaneous vehicle
attitude (/9) and incremental velocity (R, _', and _) with
respect to the established coordinate system. The Launch Vehicle
Digital Computer/Launch Vehicle Data Adapter (LVDC/LVDA)
system uses the velocity output from the platform system duringthe burn mode of flight to determine the vehicle's present velocity,
position, and acceleration. The flight program calculates the desired
vehicle pitch and yaw attitude angles as functions of time duringS-IB burn. During S-IVB burn, an optimal path to the orbital
insertion point is computed using an iterative guidance scheme.
During orbit, the vehicle's position and velocity are calculated by
using mathematical models of the earth's gravitational field, the
atmospheric drag, and the vehicle propellant vent. The orbital
guidance mode determines the guidance commands as a function
of position, prestored attitude, and time since S-IVB cutoff. Vehicleattitude control is accomplished during all periods by comparing
the present vehicle attitude (0 obtained from the platform system)to the desired attitude (X) computed by the flight program. The
difference is fed to the flight control computer as an attitude error
signal (_). The flight control computer sums the vehicle attitudeerror information with vehicle attitude rate (_) information from
the EDS control rate gyros and vehicle lateral acceleration ("r')
information (used during S-IB burn only) from the control ac-celerometers. The result of this summation is an appropriate steering
command output (fl,) that changes the vehicle's thrust vector by
either gimballing the S-IB or S-IVB engines or activating the S-IVB
auxiliary propulsion system. Figure 6-11 lists the N, G, & C system
components. Figure 6-12 is a block diagram of the Saturn IB
navigation, guidance, and control subsystem.
NAVIGATION SCHEME.
Powered Flight.
The basic navigation scheme is shown in figure 6-13, Gimbal
resolvers supply platform gimbal angles in analog form to the
LVDA. An analog-to-digital converter in the LVDA converts the
signals to the digital format required by the LVDC.
Platform integrating accelerometers sense acceleration components
and mechanically integrate them into velocity. The LVDA processes
the analog velocity data for use in the navigation equation. Within
the LVDC flight program, initial velocity imparted by the spinning
earth, gravitational velocity, and the platform velocities are alge-
braically summed. This vehicle velocity is then integrated to deter-
mine vehicle position.
Orbital Flight.
During orbital coast flight, the navigational program continually
computes the vehicle position, velocity, and acceleration from equa-tions of motion initialized at beginning of the orbit mode. In orbit,
navigation and guidance information in the LVDC can be updated
via the digital command system.
Additional navigational computations are used in maintaining
vehicle attitude during orbit. These computations establish a local
vertical which is used as a reference for attitude control.
GUIDANCE COMPUTATIONS.
The guidance function of the launch vehicle is accomplished by
computing the necessary flight maneuvers to meet the desired endconditions of the flight (e.g., inserting the spacecraft into the desired
trajectory). Guidance computations are performed within theLVDC by programmed guidance equations, which use navigation
data and mission sequence indicators as their inputs. These compu-
6-13
SectionVIInstrumentUnit
VEHICLE MOTION
ST- 124M-3
PLATFORM
LAUNCH VEHICLE
DIGITAL COMPUTERVEHICLE ATTITUDE (eY eR ep)
XI
VELOqTV(xo YI z,:
_" ="_-e, LAUNCH VEHICLE
;_i_?_; DATAADAPTER
LOCITY (X I YI Zl)
_'_._ EDS/CONTROL
._ RATE GVRO
ACCELEROMETER SIGNAL
CONDITIONER
,_'_; _' ' _ vAw LATERAL ACCELERATION FLIGHT CONTROL:_,_/__:. 'L-. I _'_'iIB_URNoNLY)
COMPUTER
YAW ACCELEROMETER _ _: -x_
CONTROL SIGNAL
PROCESSOR
PITCH ACCELEROMETER
Notes
1. VEHICLE TILTS OVER POSITION I
(HEADS DOWN),
2. ENGINE ACTUATOR LAYOUTS SHOWN
AS VIEWED FROM AFT END OF VEHICLE.
3. DIRECTIONS AND POLARITIES SHOWN
ARE TYPICAL FOR ALL STAGES.
4. +_ INDICATES ENGINE DEFLECTION
REQUIRED TO CORRECT FOR POSITIVE
VEHICLE MOVEMENT.
5. CG = CENTER OF GRAVITY
F = NOZZLE ON
EXT = ACTUATOR EXTENDED
RET = ACTUATOR RETRACTED
I_ = THRUST VECTOR ANGULAR
DEFLECTION
APS = AUXILIARY PROPULSION SYSTEM
6. ACTUATOR MOVEMENTS NOTED IN
THE POLARITY TABLE ARE
CHARACTERISTIC FOR ALL INPUT
SIGNALS OF LIKE POLARITY,
7. ALL SIGNAL ARROWS INDICATE
POSITIVE VEHICLE MOVEMENTS.
S-IB CONTROL SIGNAL POLARITY TABLE
SIGNAL & ACTUATORACTUATOR MOVEMENT
NO. +*R +_ y +_'p
I-y RET RET
I-P EXT RET
2-Y EXT RET
2-P RET EXT
3--y RET EXT
3-P EXT EXT
4-y EXT EXT
4-P RET RET
6-14
YAW OR
Z AXIS
AUXILIARY PROPULSION SYSTEM COMMANDS
INSTRUMENT _ i_
UNIT _!_
STAGE
APS RELAY PACKAGE APS RELAY PACKAGE
IIIp IIIii IIIIv Ip Ill IIV
+,p -;'y
+_F
S-IVB ACTUATORAND
0_" +_y APS NOZZLE LAYOUT
+_y -'_p IIIp
STAGE .,.II @lVp
111 [_ IIV
S-IB ACTUATOR LAYOUT Ip
Figure 6-g
S-IVB POLARITY TABLE
ACTUATOR SIGNAL & ACTION
NO. +_R "?R +_'Y +_P
]-Y EXT
I-P RET
ENGINE NO,
IIV F
Ip
III F
IIIII F
IIIp
IIIIv F
CONDITIONS DURING COAST
+_R -_R +_Y _y +tp -_p
liv F FIp F
lil F F
IIIII F F
lllp F
IIIIv F F
Section VI Instrument Unit
NAVIGATION, GUIDANCE
AND ATTITUDE CONTROL DEFINITION
FUNCTION
NAVIGATION
GUIDANCE
CONTROL
DEFINITION
DETERMINATION OF VEHICLE
POSITION, AND VELOCITY
FROM MEASUREMENTS MADE
ON BOARD THE VEHICLE.
COMPUTATION OF MANEUV_S
NECESSARY TO ACHIEVE THE
DESIRED END CONDITIONS OF
A TRAJECTORY.
EXECUTION OF NECESSARy
MANEUVERS (DETERMINED BY
THE GUIDANCE SCHEME) BY
CONTROLLING THE PROPER
HARDWARE.
ASSOCIATED EQUIPMENT
ACCELEROMETER READINGS
FROM THE ST- 12RM-3
pLATFORM.
LVDC,/LVDA.
LVDC FLIGHT PROGRAM.
LVDC FLIGHT PROGRAM.
GIMBAL ANGLE READING
FROM THE ST-124M-3
PLATFORM.
LVDC/LVDA.
LVDC FLIGHT PROGRAM.
FLIGHT CONTROL COMPUTER.
EDS/CONTROL RATE GYROS.
CONTROL ACCELEROMETERS
(IB BURN ONLY).
S-IR AND S-IVB ACTUATORS.
S-IVB AUXILIARY PROPULSION
SYSTEM.
C-H 1/,345
Figure 6-10
tations are actually a logical progression with a guidance commandas their solution. After the desired attitude has been determined
by the "optimal path" program, the guidance command is used
in the following control equation: X - 0 = q_ (figure 6-14) where:
× is the desired attitude (guidance command)
0 is the vehicle attitude
is the attitude error cbmmand
INERTIAL
ST- 124M-3 INERTIAL
PLATFORM ASSEMBLY
PLATFORM ELECTRONICS
ASSEMBLY (PEA)
PLATFORM AC POWER
SUPPLY
ACCELEROMETER SIGNALCONDITIONER
56 VDC POWER SUPPLY
DIGITAL
LAUNCH VEHICLE
DIGITAL COMPUTER
(LVDC)
LAUNCH VEHICLEDATA ADAPTER
(LVOA)
ATTITUDE CONTROL
FLIGHT CONTROL COMPUTER
(FCC)
EDS/CONTROL RATE GYROS
(EDS/CRG)
CONTROL SIGNAL
PROCESSOR (CSP)
CONTROL ACCELEROMETERS
(C/A)
S-IB CONTROl. ENGINE
ACTUATORS
_-IVB ENGINE ACTUATGRS
_UXILIARY PROPULSION
SYSTEM (AFS)
Figure 6-11
¢-N 14346
CONTROL SUBSYSTEM.
The control subsystem (figure 6-15) is designed to control and
maintain vehicle attitude by forming the steering commands used
to control the engines of the active stage.
Vehicle attitude is achieved by gimbaling the four outboard engines
of the S-IB stage or the single engine of the S-IVB stage. These
ATTITUDE ANGLES
VELOCITY
TI IU COMMAND
RECEIVER UP-DATING
& INFORMATION
DECODER
_ A'FTITUDE
CORRECTION
COMMAND
FLIGHT SEQUENCE
_'.OMMANDS
ATTITUDE CONTROL SIGNAL
FROM SPACECRAFT
1!
CONTROL' = COMPUTER
L PITCHCONTROL
ACCELEROMETER
J CO NT ROL .
ICOMMA_D
YAW JCONTROL
ACCELEROMETER
DURING S-IB BURN ONLY
s,BSTAGE
SW,TC. ELECTORS-IVB STAGE
SWITCH SELECTOR
,uSWITCH SELECTOR
S-,,STAGE
ENGINE ACTUATORS
GINES
S-IVB STAGE
ENGINE ACTUATORS
PROPULSION SYSTEM TO NOZZLES
I, CONTROLSENSORS
, TO STAGE CIRCUITRY
Figure 6-12
C-H 20019
6-15
Section Vl Instrument Unit
NAVIGATION SCHEME
ST-124M-3
FRAME_ _/_"_--w_, RESOLVER
LVDA
i ANALOG
_-- TO
DIGITAL
CONVERTER
.._ SIGNALPROCESSOR
LVDC
ATTITUDE
VELOCITY
GUIDANCE
PROGRAM
INITIAL VELOCITY
XoYoZo
INTEGRATING I'_III-'VEL.OCITY'_I_/_'VE.LO.CI.TY_I_D'I / dt
ACCELEROM ERSI I X,Y'Z')yr'×sYsZE)L:__J y1 / INITIALI _ POSITIONI GRAVlTATIO NAL
VELOCITY (X°Y°Z°)
I ORAV,TAT,ONALiGRAVITY I
j f dt liIII_ACCELERATION'"_t COMPUTATIONS I
I
A..C.C.E.L.ERAT IO N _ F/M }_(XiY I ZI ) COMPUTATIO N
TI ME -_
iVELOCITY _
PO SITIO N _._D,r (XsYsZs)
F/M
GUIDANCE
COMPUTATIO NS
IREQUIRED
ENGINE
CUTOFF
ENGINE_IGNfflON
ATTITUDE ANGLES
(XxXyXz)
1
I ATTITUDE IATTITUDECORRECTION _m" COMMAN DS
COMPUTATIONS I( *P¢'v _R)I
tPtJ_TFORM
GIMBALANGLES(8 x ey $Z)
LEGEND
dr INTEGRATION
d/tit DIFFERENTIATION
(_ ADDITIONC-H 20020
Figure 6-13
engines are gimbaled by hydraulic actuators. Roll attitude control
on the S-IVB stage cannot, of course, be controlled with a single
engine. Therefore, roll control of the S-IVB stage is accomplished
by the APS (figure 6-9). During the coast period of the mission,the S-IVB APS is used to control the vehicle attitude in all three
axes.
The control system accepts guidance commands from the guidance
system. These commands, which are actually attitude error signals,are then combined with measured data from the various control
sensors. The resultant output is the command signal to the various
engine actuators and APS nozzles.
The final computations (analog) are performed within the flight
control computer. This computer is also the central switching point
for command signals. From this point, the signals are routed to
their associated active stages and to the appropriate attitude controldevices.
NAVIGATION AND GUIDANCE COMPONENTS.
ST-124M-3 Inertial Platform Assembly.
The gimbal configuration of the ST-124M-3 offers unlimited free-
6-16
SectionVl Instrument Unit
dom about the X & Y axes, but is limited to ___45 degrees about
its Z axis (vehicle yaw at launch). See figure 6-16.
The gimbal system allows the inertial gimbal rotational freedom.
Three single-degree-of-freedom gyroscopes have their input axes
aligned along an orthogonal inertial coordinates system: X 1, Yz,
and Z 1 of the inertial gimbal. A signal generator, which is fixed
to the output axis of each gyro, generates electrical signals propor-
tional to torque disturbances. These signals are conditioned by the
servo electronics and terminate in the gimbal pivot servotorque
motors. The servo loops maintain the inner gimbal rotationally
fixed in inertial space.
The inner gimbal has three, pendulous, integrating, gyroscopicaccelerometers, oriented along the inertial coordinates X1, Yz and
Za. Each accelerometer measuring head contains a pendulous,
single-degree-of-freedom gyro. The speed of rotation of the mea-
suring head is a measure of acceleration along the input axis ofthe accelerometer. Since acceleration causes the accelerometer shaft
to be displaced as a function of time, the shaft position (with respect
to a zero reference) is proportional to velocity, and the accelero-
meter is referred to as an integrating accelerometer.
Vehicle attitude is measured with respect to the inertial platform,
using dual speed (32:1) resolvers located at the gimbal pivot points.
The outputs of these angle encoders are converted into a digitalcount in the LVDA.
During prelaunch, the ST-124M-3 platform is held aligned to the
local vertical by a set of gas bearing leveling pendulums. The
pendulum output is amplified in the platform, and then transmitted
to the ground equipment alignment amplifier. The alignment am-
plifier provides a signal to the torque drive amplifier, and then
I NAVIGATIONAL IPROGRAM
ITRUE VELOCITY (_') PRESENT POSITION (r)
/
BEST PATH PROGRAM J
/TRUE VELOCITY + PRESENT
REAL _t_ POSITION + REAL TiME +TIME PROJECTED INFLUENCES *=
DETERMINANTS FOR BEST
PATH COMPUTATION
IDESIRED ATTITUDE
(x X Xy Xz)
1
PRESENTt JVEHICLE DESIRED ATTITUDE - PRESENT
ATTITUDE _ ATTITUDE = NEEDED CHANGE
( # X 8y 87) IN ATTITUDE
ATTITU[JE ERROR
( @X ¢Y @Z )
I
I TRANSFORM TO VEHICLE J
REFERENCE FRAME
(PITCH, ROLL, & YAW)
IATTITUDE ERROR
( _'P_R '_Y)
1TO FLIGHT CONTROL SYSTEM
*e.g., POSITION OF DESTINATION
FLIGHTCONTROLSYSTEMBLOCKDIAGRAM
Figure 6-14
C-H 20021
to the platform gyro torque generator. The vertical alignment
system levels the platform to an accuracy of _2.5 arc seconds.
The azimuth alignment is accomplished by means of a theodolite
on the ground and two prisms on the platform; one fixed andone servo driven. The theodolite maintains the azimuth orientation
of the movable prism, and the computer computes a mission azi-
muth and programs the inner gimbal to that azimuth. The laying
system has an accuracy of +__5 arc seconds.
At approximately liftoff minus 17 seconds, the platform is releasedfrom an earth reference to maintain an inertial reference initiated
at the launch point. At this time, the LVDC begins navigation,
using velocity accumulations derived from the ST-124M-3 inertial
platform.
Platform Electronic Assembly (PEA).
The PEA contains the following circuitry:
a. Amplifiers, modulators, and stabilization networks for the plat-
form gimbal and accelerometer servo loops
b. Relay logic for signal and power control
c. Amplifiers for the gyro and accelerometer pick-off coil excitation
d. Automatic checkout selection and test circuitry for servo loops
e. Interlocks for the heaters and gas supply circuits.
ST-124M-3 AC Power Supply.
The ST-124-M3 platform ac power supply furnishes the power
required to run the gyro rotors and provides excitation for the
platform gimbal synchros. It is also the frequency source for theresolver chain references and for gyro and accelerometer servo
systems carrier.
The supply produces a three-phase (sine wave) output which isfixed at 26.5 V (rms) line-to-line at a frequency of 400 Hz. Three
single-phase, 20-V reference outputs (square wave) of 4.8 kHz, 1.92kHz, and 1.6 kHz are also provided. With a normal input voltage
of28 Vdc, the supply is capable of producing a continuous 250-VA
output.
INSTRUMENT UNIT
CO NTRO L-E DS
RATE GYROS
[ LVOC,LVDAFLIGHT
CONTROL
COMPUTER
Ii
AUXILIARY
PROPULSION
SYSTEM
S-IB STAGEBc
J S°IBACTUATORS
I CONTROL SIGNAL_, PROCESSOR
Jo.
, iI IACCELE ROMETERS
(S-IB BURN ONLY)
._ S-IVB JACTUATORS
Figure 6-15
C-H 20022
6-17
Section VI Instrument Unit
PLATFORM GIMBAL CONFIGURATION
FLIGHT
PATH
RESOLVER
OUTER
GIMBAL
+X
GYROIA
NCH
VERTICAL
ALIGNMENT
PENDULUMS
, INERTIAL
GIMBAL
MOTOR
÷Z j
VEHICLE
FRAME
INTEGRATING
ACCELEROMETERS
PRISM
FIXED
PRISM
SERVOTO RQUE
MOTOR
MIDDLE
SE RVOTORQUE
/V_TOR
Figure 6-16
C-H 20023
Accelerometer Signal Conditioner.
The acceterometer signal conditioner accepts the velocity signals
from the accelerometer optical encoders and shapes them before
they are passed on to the LVDA/LVDC. Additional outputs are
provided for telemetry and ground checkout.
LV Digital Computer and LV Data Adapter.
The LVDC and LVDA form an electronic digital computer
system. The LVDC is a relatively high-speed computer with the
LVDA serving as its input/output device. Any signal to or from
the computer is routed through the LVDA. See figure 6-17 and
6-18 for LVDC and LVDA characteristics.
6-18
The LVDA and LVDC are involved in four main operations:
a. Prelaunch checkout
b. Navigation and guidance computations
c. Vehicle sequencing
d. Orbital checkout
The LVDC is a general purpose computer which processes data
under control of a stored program. Data is processed serially in
two arithmetic functional areas which can, if so programmed,
operate concurrently. Addition, subtraction, and logical extractions
are performed in one arithmetic functional area while multiplication
and division are performed in the other.
The principal storage device is a random access, ferrite-core memorywith separate controls for data and instruction addressing. The
memory can be operated in either a simplex or duplex mode. In
duplex operation, memory modules are operated in pairs with the
same data being stored in each module. Readout errors in one
module are corrected by using data from its mate to restore the
defective location. In simplex operation, each module contains
different data, which doubles the capacity of the memory. However,
simplex operation decreases the reliability of the LVDC, because
the ability to correct readout errors is sacrificed. The memory
operation mode is program controlled. Temporary storage is pro-
vided by static registers, composed of latches, and by shift registers,
composed of delay lines and latches.
Computer reliability is increased within the logic sections by the
use of triple modular redundancy. Within this redundancy scheme,
three separate logic paths are voted upon to correct any errors
which develop.
CONTROL SUBSYSTEM COMPONENTS.
The control subsystem for the Saturn IB vehicles is composedof the control accelerometers in addition to the control/EDS rate
gyros, the control signal processor, and the flight control computer.
Control Accelerometers.
The body mounted control accelerometers provide lateral acceler-
ation measurements along the vehicle pitch and yaw axes. This
includes sensing the tangential component of rotational acceleration
about the pitch and yaw axes to help minimize the angle of attack
during the early portion of S-IB Burn.
The S-IB control law for the thrust vector deflection angle (_) is:
where Y is the lateral acceleration along the pitch or yaw axis
as designated, and g2 is the corresponding gain factor. The control
accelerometers sensitive axes are perpendicular to the vehicle lon-
gitudinal axis.
The lateral acceleration control is used during S-IB stage propulsion
to reduce structural loads from aerodynamic forces and to provideminimum-drift control.
Control/EDS Rate Gyros.
The vehicle angular rate error signals (_p, ¢_,, _) are supplied
by the control/EDS rate gyro package through the control signal
processor. The outputs are provided as redundant signals and arethe attitude angular rates of the vehicle about it's pitch, roll, and
yaw axes.
The angular rate error, signals control the rate of response of thevehicle to the Beta command and eliminate overshoot at the end
of the correction (reference the A1 _ term given in the control
formula in the previous paragraph on Control Accelerometers).
Control Signal Processor.
The control signal processor demodulates the ac signals from the
control-EDS rate gyros into dc signals, required by the. flight con-
trol computer. The control signal processor compares the output
signals from the triple redundant gyros and selects one each of
the pitch, yaw, and roll signals for the flight control computer.
The control signal processor supplies the control-EDS rate gyro
package with the necessary control and reference voltages. EDS
and DDAS rate gyro monitoring signals also originate within the
control signal processor, thus accounting for the EDS portion of
the control-EDS rate gyro name.
Flight Control Computer.
The flight control computer is an analog computer which converts
Section VI Instrument Unit
LVDC CHARACTERISTICS
ITEM DESCRIPTION
TYPE
MEMORY
SPEED
WORD MAKE-UP
PROGRAMMING
TIMING
INPUT/OUTPUT
GENERAL PURPOSE, DIGITAL,
STORED PROGRAM
RANDOM ACCESS, PERRITE
('rORODIAL) CORE, WITH ACAPACITY OF 32,768 WORDS OF
28 BITS EACH.
\ SERIAL PROCESSING AT 512,000
BITS PER SECONDS
MEMORY = 28 BITS, INCLUDING 2
PARITY BITS
DATA = 26 BITS
INSTRUCTION = 13 BITS
18 INSTRUCTION CODES
10 ARITHMETIC
6 PROGRAM CONTROL
1 INPUT/OUTPUT
1 STORE
COMPUTER CYCLE = 82.02 _ SEC.
BIT TIME = 1.95 p SEC.
CLOCK TIME = 0.49 _ SEC.
EXTERNAL, PROGRAM CONTROLLED
Figure 6-17
C-H 20024
ITEM DESCRIPTION
INPUT/OUTPUTRATE
SWITCH
SELECTOR
TELEMETRY
COMMAND
RECEIVER
DATA
TRANSMITTER
COMPUTER
INTERFACE
UNIT
DELAY LINES
OUTPUT TO
LAUNCH COM-
PUTER
INPUT FROM
RCA-110 GCC
SERIAL PROCESSING AT 512,000
BITS PER SECOND
8 BIT INPUT
15 BIT OUTPUT
14 BITS FOR INPUT DATA
38 DATA AND IDENTIFICATION
8ITS PLUS VALIDITY BIT AND
PARITY BIT
15 BiTS ADDRESS PLUS I DATA
REQUEST BIT
10 SITS FOR INPUT DATA PLUS I
BIT FOR DATA READY INTERRUPT
3 FOUR-CHANNEL DELAy LINES
FOR NORMAL OPERATION
I FOUR-CHANNEL DELAY LINE
FOR TELEMETRY OPERATIONS
41 DATA AND IDENTIFICATION
BITS PLUS DISCRETE OUTPUTS
14 BITS FOR DATA PLUS
INTERRUPT
C-H 20025
Figure 6-18
attitude correction commands (,tt,), angular change (_), and vehicle
lateral acceleration ( _; ) information (during S-IB burn mode only)
into APS thruster nozzle and/or engine actuator positioning com-
mands.
Input signals to the flight control computer include:
a. Attitude correction commands (,t/,) from the LVDC/LVDA or
spacecraft.
b. Angular rates (_) from the control-EDS rate gyro package, via
the control signal processor.
6-19
Section VI Instrument Unit
c. Vehicle lateral accelerations (_;) from the IU-mounted control
accelerometers (during S-IB burn only).
Output signals from the flight control computer include:
a. Command signals to the engine actuators (tic)
b. Command signals to the APS thruster nozzles (tic)
c. Telemetry outputs which monitor internal operations and func-tions.
FLIGHT PROGRAM.
The flight program which is structured modularly, is composedof two basic subsystems: the control subsystem and the application
subsystem. The control subsystem controls the sequence and order
of execution of all programmed functions in the application subsys-tems.
The control subsystem consists of the control program and a com-
mon communications area for inter-module communications, com-
mon data and indicators, and mission or vehicle dependent para-
meters. The control program, composed of sub-programs and tables,
controls the execution of application modules, services interrupts,
and routes control to the appropriate application module on apriority basis, and provides utility operations.
The application subsystem provides a master pool of all defined
LVDC functions which can be used to generate a total flight
program configuration for any given mission task. It is composedof a collection of relatively independent, closed program modules.
Each application module is designed to perform one or a number
of related functions. These functions include navigation, guidance,
attitude control, event sequencing, redundancy and data manage-
ment, ground command processing, and hardware evaluation.
For purposes of discussion, the flight program is divided into five
subelements: the powered flight major loop, the orbital flight pro-
gram, the minor loop, interrupts, and telemetry. There is also an
LVDC pre-flight program which supports launch vehicle checkout
and performs the functions that prepare the LVDC for enteringthe flight mode.
Orelaunch and Initialization.
Until just minutes before launch, the LVDC is under control of
the ground control computer (GCC). At approximately T-10 min-
utes, the GCC issues a prepare-to-launch (PTL) command to the
LVDC. The PTL routine performs the following functions:
a. Monitors accelerometer inputs, calculates the platform-off-level
indicators, and telemeters accelerometer outputs and time
b. Performs reasonableness checks on particular discrete inputs andalerts
c. Interrogates the LVDC error monitor register
d. Keeps all flight control system ladder outputs zeroed, whichkeeps the engines in a neutral position for launch
e. Processes the GRR interrupt and transfers LVDC control to
the flight program
f. Samples platform gimbal angles.
At T-22 seconds, the launch sequencer issues a GRR alert signal
to the LVDC and GCC. At T-17 seconds, a GRR interrupt signal
is sent to the LVDC and GCC. With the receipt of this signal,
the PTL routine transfers control of the LVDC to the flight program.
When the GRR interrupt is received by the LVDC, the followingevents take place:
a. The LVDC sets time base zero (To)
6-20
b. Gimbal angles and accelerometer values are sampled and storedfor use by flight program routines
c. Time and accelerometer readings are telemetered
d. All flight variables are initialized
e. The GCC is signaled that the LVDC is under control of the
flight program.
During the time period between GRR and liftoff, the LVDC begins
to perform navigational calculations and process minor loops. Atliftoff time base 1 (T1) is initiated.
Powered Flight Major Loop.
The major loop contains the navigation and guidance calculations,
timekeeping, and other repetitive operations of the flight program
that do not occur on an interrupt basis. Its various routines are
subdivided by function. Depending upon mode of operation and
time of flight, the program will follow the appropriate sequenceof routines.
The accelerometer processing routine accomplishes two main objec-
tives: it accumulates velocities as measured by the platform and
detects velocity measurement errors through "reasonableness" tests.
The boost navigation routine combines gravitational acceleration
with measured platform data to compute position and velocity.
The "pre-iterative" guidance mode, or "time-tilt" guidance pro-
gram, is that part of the flight program which performs from liftoff
until the end of the S-IB burn. The guidance commands issued
during the time-tilt phase are functions of time and engines out
only. This phase of the program is referred to as open loop guidance,
since vehicle dynamics do not affect or influence the guidancecommands. When the launch vehicle has cleared the mobile
launcher, the time-tilt program then initiates a roll maneuver to
align the vehicle with the proper azimuth and a time-tilt pitch
maneuver. Time-tilt yaw guidance commands are computed as a
tabular function of time. The roll maneuver is completed at approx-
imately T + 38 seconds while the pitch and yaw time-tilts continue
until the guidance commands are frozen (tilt arrest) prior to start
of Time Base 2 (Low Level Sense) and Inboard Engine cutoff.
Provisions are made for early engine out guidance modifications
for the first detected engine failure only. In Time Base 1, this consists
of freezing the pitch guidance command for a specified length of
time and modification of the time-tilt computations. In addition,certain other parameter adjustments are made. In Time Base 2,
the only change is in the value of the sine function in the zero
test computations for outboard engine failure. Time base 3 ('1"3)
commences where the outboard engine thrust delay is sensed fol-lowing cutoff.
The iterative guidance mode (IGM) routine, or "path adaptive"
guidance, commences approximately 35 seconds after S-1VB stageignition, and continues until the end of the S-IVB burn. Cutoff
is commanded when the velocity required for the target orbit has
been reached. IGM employs optimizing techniques, based on the
calculus of variations, to determine a minimal propellant flightpath which satisfies mission requirements. Since the IGM reacts
to vehicle dynamics, it is referred to as closed loop guidance.
Orbital Flight Program.
The orbital flight routines consist of an executive routine, telemetrytime-sharing routines to be employed while the vehicle is over
receiving stations, navigation, guidance, an d timekeeping computa-tions.
When in orbital mode, the flight program will process telemetry
acquisition and loss determination once per eight seconds, event
sequencing and interrupt processing as required, in addition to
orbitalnavigation,guidance,andcontrol.Otherfunctionsincludeminorloop,tenpersecond;minorloopsupportanddiscretepro-cessing,oncepersecond;andgimbalanglereadfororbitalnaviga-tion,onceeveryfourseconds(orbitalnavigationparametersarecalculatedoncepereightseconds).Minor Loop.
The minor loop contains control system computations. Since the
minor loop is used for vehicle control, minor loop computations
are executed at the rate of 25 times per second during the powered
phase of flight. However, in earth orbit, a rate of only ten executionsper second is required for satisfactory vehicle control. Rate limiting
of the output commands prevents the flight control systems frommaneuvering the LV at rates that exceed safe limits.
The supporting control functions of computing attitude change
increments and coefficients for gimbal-to-body transformation re-
quired for attitude error command computations are called minor
loop support functions. These functions must be performed once
per computation cycle during boost and once per second duringorbit.
Interrupts.
An interrupt routine permits interruption of the normal program
operation to free the LVDC for priority work, and may occur at
any time within the program sequence. When an interrupt occurs,
the interrupt transfers LVDC control to a special subroutine which
identifies the interrupt source, performs the necessary subroutines,
and then returns to the point in the program where the interrupt
occurred. Figure 6-19 lists the LVDC interrupts.
Telemetry Routine.
A programmed telemetry feature is also provided as a method
of monitoring LVDC and LVDA operations. The telemetry routine
transmits specified information and data to the ground via IU
telemetry equipment. In orbit, telemetry data must be stored at
times when the vehicle is not within range of a ground receiving
station. This operation is referred to as data compression. The storeddata is transmitted on a time-shared basis with real-time telemetry
when the LV is within range of a station.
LVDC INTERRIIPTS
DISCRETE BACKUPS.
Certain events, are particularly important to the flight program
since they are time base references. These events are indicated
to the flight program by primary and backup signal paths. The
primary signals are recognized as program interrupts; the backup
signals are treated as periodically serviced discretes.
Because switch selector commands are functions of time (relative
to one of the time bases); accurately timed switch selector com-mands could not be generated if one of the time-base-initiated
discrete signals were missed.
The execution time for any given major loop, including minor loop
computations and interrupts, is not fixed because the number of
modules executed is dependent on the flight mode, the discrete
and interrupt processing requirement is not predictable, and the
number of minor loops in each major loop is not fixed.
MODE AND SEQUENCE CONTROL.
Mode and sequence control involves most of the electrical/elec-
tronic systems in the launch vehicle. However, in this section, the
discussion will deal mainly with the switch selectors and associated
circuitry.
The LVDC memory contains a predetermined number of sets of
instructions which, When initiated, induce portions of the launch
vehicle electrical/electronic systems to operate in a particular mode.
Section VI Instrument Unit
INTERRUPT
STORAGE
REGISTER BIT
7
8b
8a
9
10
11
12
LVDC DATA
WORD BIT
POSITION
11
10
9
8
7
6
5
4
4
3
2
1
S;gn }
FUNCTION
RCA-110A INTERRUPT
* S-IB LOW LEVEL SENSORS DRY "A"
RCA-110A INTERRUPT
S-IVB ENGINE OUT "B"
* S-IBOUTBOARD ENGINES
CUTOFF "A"
. MANUAL INITIATION OF
S-IVB ENGINE
CUTOFF "A"
GUIDANCE REfeRENCE RELEASE
COMMAND DECODER INTERRUPT
"B"
COMMAND DECODER INTERRUPT
"A"
SIMULTANEOUS MEMORY ERROR
SPARE
• INTERNALTO THE LVDC
C-H 20026
*TIMES ARE MISSION DEPENDENT
. Figure 6-19
Each mode consists of a predetermined sequence of events. The
LVDC also generates appropriate discrete signals such as engine
ignition, engine cutoff, and stage separation.
Most selection and initiation can be accomplished by an automatic
LVDC internal command, an external command from ground
checkout equipment or IU command system, or by the flight crew
in the spacecraft.
The flexibility of the mode and sequence control scheme is suchthat no hardware modification is required for mode and flight
sequence changes. The changes are accomplished by changing the
instructions and data in the LVDC memory.
Switch Selector.
Many of the sequential operations controlled by the LVDC are
performed through a switch selector located in each stage. Theswitch selector decodes digital flight sequence commands from the
LVDA/LVDC and activates the proper relays, either in the units
affected or in the stage sequencer.
Each switch selector can activate, one at a time, up to 112 different
relays in its stage. The selection of a particular stage switch selector
is accomplished through the command code. Coding of flight se-
quence commands and decoding by the stage switch selectors re-
duces the number of interface lines between stages and increases
the flexibility of the system with respect to timing and sequence.
In the launch vehicle, which contains three switch selectors, up
to 336 different functions can be controlled, using only 28 lines
from the LVDA. Flight sequence commands may be issued at timeintervals as short as 100 milliseconds.
To maintain power isolation between vehicle stages, the switch
selectors are divided into sections. The input sections (relay circuits)
of each switch selector receive their power from the IU. The output
sections (decoding circuitry and drivers) receive their power from
the stage in which the switch selector is located. The inputs and
outputs are coupled together through a diode matrix. This matrix
decodes the 8-bit input code, and activates a transistorized output
driver, thus producing a switch selector output.
The output signals of the LVDA switch selector register, with the
exception of the 8-bit command, are sampled at the control distrib-
utor in the IU and sent to IU PCM telemetry. Each switch selector
also provides three outputs to the telemetry system within its stage.
The switch selector is designed to execute flight sequence commands
6-21
Section Vl Instrument Unit
given by the 8-bit code or by its complement. This feature increases
reliability and permits operation of the system, despite certain
failures in the LVDA switch selector register, line drivers, interface
cabling, or switch selector relays.
The flight sequence commands are stored in the LVDC memory,
and are issued by the flight program. When a programmed input/
output instruction is given, the LVDC loads the 15-bit switch
selector register with the computer data.
The switch selector register, bits I through 8, represents the flight
sequence command. Bits 9 through 13 select the switch selector
to be activated. Bit 14 resets all the relays in the switch selectors
in the event data transfer is incorrect, as indicated by verification
fault information received by the LVDA. Bit 15 activates the ad-
dressed switch selector for execution of the command. The switch
selector register is loaded in two passes by the LVDC: bits 1 through
13 on the first pass, and either bit 14 or bit 15 on the second
pass, depending on the feedback code. The LVDA/LVDC receives
the complement of the code after the flight sequence command
(bits 1 through 8) has been picked up by the input relays of the
switch selector. The feedback (verification information) is returned
to the LVDA, and compared with the original code in the LVDC.
If the feedback agrees, the LVDC/LVDA sends a read command
to the switch selector. If the verification is not correct, a reset
SWITCH SELECTORFUNCTIONAL CONFIGURATIONNOTES: SIGNAL RETURN LINES FROM THE SWITCH SELECTORS, THROUGH THE CONTROL
DISTRIBUTOR, TO THE LVDA ARE NOT SHOWN IN THIS FIGURE.
THE LETTERS USED TO LABEL INTERSTAGE CONNECTIONS BETWEEN UNITS ARE
NOT ACTUAL PIN OR CABLE CONNECTORS. THE LETTER CODE IS DENOTEDBELOW:
a = 8-DIGIT COMMAND (8 LINES)
b = FORCE RESET (REGISTER) (I LINE + 1 REDUNDANT LINE)e = REGISTER VERIFICATION (8 LINES)
d = READ COMMAND (1 LINE + 1 REDUNDANT LINE)
; I STAOESELECTL,NESg __ (1 LINE + ] REDUNDANT LINE)
j = b, c, d, e, f, g, AND h to IU TELEMETRY
k = REGISTER TEST )
ZERO INDICATE ] TO STAGE TELEMETRYSW SEL OUTPUT I (I LINE EACH)
m = +28 VDC FROM THE INSTRUMENT UNIT
GROUND
COMPUTERilltiilil iabed F gh
IsSELECTOt
TEST
PANEL 7
m mm mm m m mm_ ml a mmmmmm me mini nm mmmmlmum m | m m am lmmllm
i' fitU
LVDA _ f 5DIGITAL g ZZ]
AUXILIARY
ROWER
DISTRIBUTOR
AUXILIARY
ROWER
DISTRIBUTOR
112 OUTPUT LINES SWITCHSELECTOR
| m mm mm mmmaum mmmmmm mill Imml a mmm mmmmm|
STAGE TELEMETRY
AND DDAS
I STAGE J_ STAGE + 28 VDC J-_J SWITCH
SELECTORCONTROL 112 OUTPUT LINES /CIRCUITRY /
iIm | i mn ii ummmluuu im m m
TELEMETRY J
AND DDAS |
i_i i I'i / I I J coNTROLizJ _ T i I I I I DISTRIBUTOR
II]"]/ll17.1 I I I TJ.I | Ju .
I=l , I I I T 1 l+ 28 VDC
,I--I I I I I ITln o-,,_ iii iillTro a bd f hg ern
I., ,L,IJ, JI
rM I
IU
S-IVB STAGEli m mmimm m lmmommamammm m m mummm m m | mm | mmmmm m lm I| mmmlmm a nil lmmmmlmnlmmm
AND DDAS _ b
J J_ STAGE + 28 VDC/-_ SWITCHSTAGE 112 OUTPUT LINES / SELECTORCONTROL
CIRCUITRY
S-IS STAGEmmmmmm tm mnmimnmm m mmmm m iNm nl m m m m emmm mlmlmmm m m i m Immmmm m t
c_ 2oo27
Figure 6-20
6-22
Section VI Instrument Unit
LVDCSWITCHSELECTORINTERCONNECTIONDIAGRAM
LVDC
15- BIT SERIAL
Mf
LVDA J STAGE SELECT I
1121, I l° I,t I '31
CONTROL
DISTRIBUTOR
FLIGHT JCOMMAND
1 THROUGH 8 I
8 LINES /
| t
RES I I1, I
SIGNAL
RETURN
IU 6D COM
; i,i, ,IJ, ,l,,i
• A
DIGITAL INPUT
MULTIPLEXER
I 1 THROUGH 8
1 b
I e-BIT SERIAL
lU
SWITCH
SELECTOR
I S-IVB
SWITCH
SELECTOF
S-IB
SWITCH
SELECTO_
C-H 20028
Figure 6-21
command is give n (forced reset), and the LVDC/LVDA reissues
the 8-bit command in complement form.
Figure 6-20 illustrates the Saturn IB switch selector functional
configuration. All switch selector control lines are connectedthrough the control distributor in the IU to the LVDC and the
electrical support equipment.
The LVDC switch selector interconnection diagram is shown in
figure 6-21. All connections between the LVDA and the switch
selectors, with the exception of the stage select inputs, are connected
in parallel.
Operation Sequence.
The Saturn IB operationsequence starts during the prelaunch phase
at approximately T-14 hours, when the electrical power from the
ground support equipment is applied to all stages of the launch
vehicle. During this time, the sequencing is controlled from thelaunch control center/mobile launcher complex, utilizing both
manual and automatic control to check out the functions of the
entire launch vehicle. After the umbilicals are disconnected, the
sequencing is primarily controlled by the flight program withinthe LVDC.
Since flight sequencing is time-phased, the sequencing operation
is divided into four primary time bases. Each time base is related
to a particular flight event. These time bases are defined in the
following paragraphs.
Time Base No. 1 (T 0. T 1 is initiated by either of the liftoff signals
provided by deactuation of the liftoff relays in the IU at umbilicaldisconnect. Redundancy for this function is provided by three wires
through the umbilical supplying ground power to the three discon-
nect relays. At IU umbilical disconnect, interruption of any two
of these three signals will set T;. If the time since GRR is greater
than or equal to 17.4 seconds and less than 150 seconds and eitheror both liftoff signals are present, T 1 will be set, indicating liftoff
has occurred. See figure 6-22 for a logic diagram of these functions.
No "negative backup" (i.e., provisions for the LVDC to return
to prelaunch conditions) is provided because, in the event T1 began
by error, the launch vehicle could safely complete T 1 on the pad
without catastrophic results.
Time Base No. 2 (7-2). After arming the S-IB propellant level sensors
through the S-IB switch selector, the LVDC will initiate time base
No. 2 (T 2) upon receiving either of two redundant fuel or LOX
level sensor signals, if sufficient downrange velocity exists at thattime. However, if Guidartce Reference Failure (GRF) has occurred,
the LVDC will bypass the velocity test and initiate time base
No. 2.
Use of the downrange velocity reading provides a safeguard against
starting T2 on the pad, should T 1 be started without liftoff. Further-
more, if T 2 is not established, no subsequent time bases can bestarted. This ensures a safe vehicle, requiring at least one additional
failure to render the vehicle unsafe on the pad.
Time Base No. 3 (7"3). After arming (Tops grouping) the S-IB LOX
Depletion Sensors through the S-tB switch selector, the LVDC shallinitiate time base No. 3 upon receiving either of two redundant
outboard cutoff signals. The S-IB outboard engines cutoff "A'"
signal (INT 5) is the primary signal, with the S-IB outboard engines
cutoff "B'" (D123) as the backup signal, for starting T3.
Time Base No. 4 (T4). Anytime after T 3 + 10 seconds, any of
the following four combinations of events will start T4.
6-23
Section VI Instrument Unit
a. Both S-IVB engine out "A" and "B" indications from the thrust
OK pressure switch when the S-IVB engine shuts down.
b. Either engine out indication and the velocity cutoff command.
c. Either engine out indication and a velocity change indication
from the LVDC signaling a loss of thrust.
d. The velocity cutoff command and LVDC indication of loss ofthrust.
Redundant S-IVB cutoff commands are issued at the start of T4
to ensure against starting this time base with S-IVB engine thrustpresent.
MEASUREMENT AND TELEMETRY.
Measurement and telemetry instrumentation within the IU consists
of a measuring subsystem, a telemetry subsystem', and an antennasubsystem. This instrumentation monitors certain conditions and
events which take place within the IU and transmits monitored
signals to ground receiving stations. Telemetry data is used on the
ground for the following purposes:
a. Preflight checkout of the launch vehicle.
b. During vehicle flight, for immediate determination of vehicle
condition and for verification of commands received by the IUcommand system.
c. Postflight scientific analysis of the mission.
MEASUREMENTS.
The requirement for measurements of a wide variety has dictated
the use of many types of transducers at many different locations.
However, a discussion of each transducer type is beyond the scope
of this manual. The parameters measured include such things as
acceleration, angular velocity, flow rate, position, pressure, temper-
ature, voltage, current, and frequency.
Conditioning of measured signals is accomplished by amplifiers
or converters located in measuring racks. Each measuring rack has
a capacity of 20 signal conditioning modules. In addition to its
conditioning circuitry, most of the signal conditioning modules
contain the capability for simulating the transducer inputs. This
capability is used for prelaunch calibration of the signal condi-tioners.
Measurement signals are generally routed to their assigned telem-
etry channel by the measuring distributors. The measuring dis-
tributors contain the capability to connect different measurements
to the same telemetry channels during ctifferent flight periods.
Selected FM measurements are switched to digital data acquisition
system (DDAS) channels for ground checkout and then returnedto the FM link after checkout.
TELEMETRY.
The function of the IU telemetry (TM) system is to format and
transmit measurement signals. See figure 6-23 for a diagram of
TIME BASEDEFINITION
6-24
LIFTOFF A % "_"
LIFTOFF B
GRR >_17.4 SEC TI
GRR < 150 SEC
TIME BASE NO. I
PROPELLENT LEVEL SENSOR ENABLE
FUEL LEVEL
SENSOR NO. 1
LOX LEVEL
SENSOR NO. 2
FUEL LEVEL
SENSOR NO. 2
LOX LEVEL
SENSOR NO. 3 L
DOWNRANGE VELOCITY _ 500 M/S_F_._ %
Fl 9GUIDANCE REFERENCE
FAILURE INDICATION
TIME BASE NO. 2
NOTE:
[_IF PREVENTS THE LVDC FROM RECEIVING ORA FAILURE
RECOGNIZING THE S-IB OBCO SIGNALS, THE LVDC WILL
ISSUE AN OBCO COMMAND AND START T3 AT T2 + 13 SEC.
LOX DEPLETION CUTOFF ENABLE (TOPS GROUPING)
LOXDE LET,ON ______I"FUEL DEPLETION
CUTOFF ENABLE
FUEL DEPLETION
SENSOR ACTUATED
T2 + 13.0 S T3
TIME BASE NO. 3
S-IVB ENGINE OUT B
S-,VBENG,NE l/
VELOCITY CHANGE -< 1 M/SEC
FOR PAST BML
1LTIME BASE NO. 4
T4
C-H 2C_29
Figure 6-22
the IU TM system. The approximately 225 measurements takenon the IU are transmitted via two TM links as follows:
a. Link P1, Pulse Code Modulation/Frequency Modulation
(PCM/FM).
b. Link FI, Frequency Modulation/Frequency Modulation (FM/FM).
Multiplexing.
To enable the two TM links to handle the approximately 225
measurements, both frequency sharing and time sharing multiplex-
ing techniques are used.
Two Model 270 time sharing multiplexers are used in the 1U
telemetry system. Each one operates as a 30 x 120 (30 primary
channels, each sampled 120 times per second) multiplexer with
23 of the primary data channels containing provisions for being
submuhiplexed. Each sub-multiplexed primary channel forms ten
sub-channels, each sampled at 12 times per second. Twenty-seven
of the 30 primary channels are used for measurement data, while
the remaining three are used for references.
The Model 270 also has an integral calibration generator for inflight
calibration capability. Upon command from the TM calibrator,the calibration generator seeks the next available master frame,
inhibits the normal measurement data input, _nd applies a sequence
of five calibration voltages to all data channels. Each voltage level
is sustained for one master frame, thus requiring five frames or
approximately 400 milliseconds for a complete calibration sequence.
Two Model 410 remote digital multiplexers are used in the IU
TM system. The Model 410 can accept up to 100 inputs in the
form of discretes, digital data, or a combination. This data is
temporarily stored as 10-bit digital words and then transferred to
the Model 301 programmed format.
Low level conditioned analog signals are fed to subcarrier oscillators
(SCO) in the FI TM oscillator assembly (Model BI). The Model
B 1 has the capability of handling 27 continuous data channel inputs
by utilizing IRIG channels 14 and 17 for FM/FM/FM. Each input
signal is applied to a separate SCO, and each SCO produces a
different output frequency. The SCO outputs are combined and
the composite signal frequency modulates the FI RF assembly.
The PCM/FM system performs a dual function. During flight, it
serves as a TM link; during prelaunch checkout, it serves as an
interface with the digital GSE. PCM techniques provide the high
degree of accuracy required for telemetering certain signal types.
The Model 301 unit accepts and digitalizes analog inputs from
the Model 270 and serializes these signals with digital signals from
the Model 410 and d.irect discrete inputs. The Model 301 output
is a serial train of digital data, which modulates the RF assembly.
During prelaunch checkout, the Model 301 output also feeds
through coaxial cable to the GSE.
Both TM RF assemblies are VHF/FM transmitters and use combi-
nations of solid state and vacuum tube electronics. The transmitter
outputs couple into a single antennasystem containing two omnidirectional antennas.
The PCM/FM systems of the IU and S-IVB are cross-strapped
to provide a redundant transmission path for flight control andother critical measurements. This arrangement routes the data fromone IU Model 270 to the S-IVB Model 301, while the data from
one of the S-IVB Model 270's is routed to the IU Model 301.
COMMAND SYSTEM.
The IU command system is used to transmit digital information
from ground stations to the Launch Vehicle Digital Computer
Section VI Instrument Unit
(LVDC) in the IU. Figure 6-24 is a block diagram of the command
system for Saturn lB.
The commands and.data to be transmitted to the vehicle originate
in the Mission Control Center (MCC) in Houston, Texas, and are
sent to the remote ground stations of the Spaceflight Tracking and
Data Network (STDN). At the ground stations, the command
messages (in digital form) are processed and temporarily stored.
On command from MCC they are modulated onto a 450 MHz
RF carrier by a Digit.al Command Syste m (DES) and transmitted.
The signal is received and demodulated by the command receiver.
The resultant signal is fed to a command decoder where addressverification and final message decoding is accomplished. From the
command decoder, the command message (in digital form) is sent
through the Launch Vehicle Data Adapter (LVDA) to the LVDC.
Verification of the acceptance or rejection of the command message
is telemetered to the ground station via the IU telemetry system.
If a message is rejected, pertinent data concerning the rejectionwill also be telemetered.
At the ground station, the verification and acceptance data is
recovered from the telemetry message and fed to the Des through
special circuitry. The remainder of the telemetry data is forwarded
to the MCC. If an acceptance message is not received by the ground
station within a preset length of time, the same message will be
retransmitted. Presently, a total of four such attempts will be made
before the message is abandoned.
The LVDC is programmed to receive two types of command wordsfrom the command decoder: mode words and data words.
The LVDC can be programmed to recognize as many as 26 different
mode command words. Many of these command words are common
to all flights while others are programmed only for particularmissions. Common mode commands include:
a. Time base update
b. Navigational update
c. Execute switch selector routine
dl Telemeter memory contents
e. Terminate command routine.
Data words, as the name implies, contain data to supplement modecommands. The number of data words varies with the mode com-
mand involved. For example, a time base update requires only
one data word while a navigation update requires more than thirtydata words.
SATURN TRACKINGINSTRUMENTATION.
Radio tracking determines the vehicle's trajectory providing data
for mission control, range safety, and postflight evaluation of vehicle
performance.
C-BAND RADAR.
The function of the C-band radar transponder is to increase the
range and accuracy of the radar ground stations equipped with
AN/FPS-16, and AN/FPQ-6 radar systems. C-band radar stations
at the Kennedy Space Center, along the Atlantic Missile Range,
and at many other locations around the world, provide global
tracking capabilities. Two C-band radar transponders are carried
in the IU to provide radar tracking capabilities independent ofthe vehicle attitude. Two antennas (one for each transponder) are
6-25
Section VI Instrument Unit
TELEMETRYSYSTEM
ANTENNA
• 3A58!
POWER DIVIDER
601A436
COAXIALSWITCH 60|A563
. DIRECTIONAL
602A642
TM
RFCOUPLER 602A445
I= _.e
270
MUX
PCM/DDAS ASSY
1S-IVB STAGE
602A580
ANTENNA
RF
DUMMY
TM CALIB.
• CONTROL
UNIT
TM CALIBRATOR ASSY
2A644
602A566
6-26Figure 6-23
CALII_RATION SIGNAL
C-H 14360
TRANSMITTINGDIGITALEQUIPMENTH COMMANDSYSTEM
J STA_%!
C-H 20030
Figure 6-24
located 180 ° apart on the IU outer surface providing an omniantenna pattern around the vehicle.
The transponder consists of a single compact package. Major ele-
ments include an integrated RF head, an IF amplifier, a decoder,
over-interrogation protection circuitry, a fast-recovery solid-state
modulator, a magnetron, a secondary power supply, and transducers
for telemetry channels.
The transponder receives a coded double pulse interrogation from
Section VI Instrument Unit
ground stations and transmits a single-pulse reply in the same
frequency band.
Two conditioned telemetry outputs are provided to the telemetry
system: input signal level and input PRF.
The characteristics of the C-band radar transponder are given infigure 6-25.
GROUND SUPPORT EQUIPMENT.
The IU, because of its complex nature, requires the services ofmany types of GSE (mechanical, pneudraulic, electrical, electronic)
and personnel. This section of the manual is limited to a very
brief description of the IU GSE.
There are three primary interfaces between the IU and its GSE.
One is the IU access door, used during prelaunch preparations
for battery installation, ordnance servicing, servicing IU equipment
and S-IVB forward dome. The second interface is the umbilical,
through which the IU is furnished with ground power, purging
air/GN2, coolant for environmental control and hardwire links with
electrical/electronic checkout equipment. The third interface is the
optical window through which the guidance system ST-124M-3
inertial platform is aligned.
IU ACCESS DOOR.
The structure of the IU consists of three 120-degree segments of
aluminum honeycomb sandwich, joined to form a cylindrical ring.
After assembly of the IU, a door assembly provides access to the
electronic equipment .inside the structure. This access door has beendesigned to act as a load supporting part of the structure in flight.
Work platforms, lights, and air-conditioning are used inside the
1U to facilitate servicing operations. When the spacecraft is being
C-BAND TRANSPONDERCHARACTERISTICS
RECEIVER CHARACTERISTICS
FREQUENCY (TUNABLE EXTERNALLY)
BANDWIDTH (3 db)
OFF-FREQUENCy REJECTION
SENSITIVITY (99% REPLY)
MAXIMUM INPUT SIGNAL
INTERROGATION CODE
PULSE WIDTH
PULSE SPACING
TRANSMITTER CHARACTERISTICS
FREQUENCY (TUNABLE EXTERNALLY)PEAK "POWER OUTPUT
PULSE WIDTH
PULSE JITTER
PULSE RISE TIME (10% TO 90%)
DUTY CYCLE
VSWR OF LOAD
PULSE REPETITION RATE
TRANSPONDER CHARACTERISTICS
RECOVER TIME
FIXED DELAY
DELAY VARIATION WITH SIGNAL LEVEL
POWER REQUIREMENTS
PRIMARy CURRENT DRAIN
5690 ± 2 MHz
8-12 MHz
50 db IMAGE; 80 clb MI NIMUM, 0.15 TO 10,000 MHz
-65 dbm OVER ENTIRE FREQUENCY RANGE AND ALL ENVIROMENTS+20 dbm
DOUBLE PULSE
0.25 TO 1.0 _SEC (DOUBLE PULSE)8±0.1 _SEC
5765 ± 2 MHz
400 WATTS MINIMUM, 700 WATTS NOMINAL
1.0 ±0.1 _SEC
0.020 _SEC MAXIMUM FOR SIGNALS ABOVE -55 dbrn
0.1 _SEC MAXIMUM
0.002 MAXIMUM
1.5:1 MAXIMUM
10 TO 2000 pps; OVERIINTERROGATION PROTECTION ALLOWS
INTERROGATION AT MUCH HIGHER RATES WITH COUNT-DOWN;REPLIES DURING OVERINTERROGATION MEET ALL REQUIREMENTS
50 uSEC SINGLE PULSE, 62 _SEC DOUBLE PULSE MAXIMUM FOR INPUT
SIGNAL LEVELS DIFFERING BY UPTO 65 ova(RECOVERS TO PULL
SENSITIVITY WITH NO CHANGE IN TRANSMITTER REPLY POWER OR
FREQUENCy WITH MULTIPLE RADARS INTERROGATING
SIMULTANEOUSLY)
3.0±0.01 _SEC
50 NANOSECONDS MAXIMUM FROM -65 dbm TO 0 dbm
24 TO 30 VOLTS
0.7 AMPERE STANDBY; 0.9 AMPERE AT 1000 pps
Figure 6-25
C-H 20031
6-27
SectionVIInstrumentUnit
IUfSLA ALIGNMENT
TYPICAL
3 PLACE
i I
,,
i
1.020 DIA1.010
GUIDE PiN
_L 50MIN
F " "
ALL DIMENSIONS ARE IN INCHES
C-H 20032
Figure 6-26
Ig rSLA _ECHANICAL ATTACHMENT
_--- INSTRUMENT
UNIT
Figure 6-27
C-H 20033
fueled through the IU access door, a special protective cover is
installed inside the IU to protect components from any possible
volatile fuel spillage.
Approximately 20 hours before launch, the IU flight batteries, each
weighing 165 pounds, are activated in the battery shop and installed
in the IU through the access door.
At approximately L-6 hours, the service equipment is removedand the access door is secured.
6-28
SLA
IUNIT
IU (REF)
ALL DIMENSIONS ARE IN iNCHES
C-H 20034
IU UMBILICAL.
Figure 6-28
The physical link between the IU and the GSE is through the
umbilical connection, located adjacent to the access door, that mates
with the service arm umbilical. The umbilical is made up of
numerous electrical connectors, two pneudraulic couplings
and an air conditioning duct. The electrical connectors provide
ground power and the electrical/electronic signals necessary for
prelaunch checkout of the IU equipment. The quick-disconnect
couplings provide for circulation of GSE supplied Oronite coolant
fluid through the onboard IU/S-IVB ECS cooling system until
umbilical disconnect at liftoff. The air conditioning duct provides
for compartment cooling air or purging GN 2.
The service arm umbilical is retracted at liftoff and a spring loaded
door on the IU closes covering the umbilical plate.
OPTICAL ALIGNMENT.
The IU contains a window through which the ST-124M-3 inertial
platform has its alignment checked and corrected by a theodolite
located in a hut on the ground and a computer feedback loop.
By means of this loop, the launch azimuth can be monitored,
updated, and verified to a high degree of accuracy.
lU/SLA INTERFACE.
MECHANICAL INTERFACE.
The IU and spacecraft-LM adapter (SLA) are mechanically aligned
with three guide pins and brackets as shown in figure 6-26. These
pins facilitate the alignment of the close tolerance interface bolt
holes as the two units are joined during vehicle assembly. Six bolts
are installed around the circumference of the interface and sequen-
tially torqued using a special MSFC designed wrench assembly.
These six bolts secure the IU/SLA mechanical interface. (See figure6-27.)
Electrical Interface.
The electrical interface between the IU and spacecraft consists of
three 61-pin connectors. (See figure 6-28.) The definition and
function of each connector is presented in the following paragraphs.
IO/Spacecraft Interface Connector J-1.
This connector provides lines for power and for control, indication,
and EDS circuitry.
IO/Spacecraft Interface Connector J-2.
This connector provides lines for power, control, and indications
for the Q-ball circuitry and the EDS circuitry.
IU/Spacecraft Interface Connector J-3.
This connector provides lines for power and control, indication,
and EDS circuitry.
EXPERIMENTS.
M415, THERMAL CONTROL COATINGS.
The objective of Experiment M-415 is to determine the degradation
effects of prelaunch, launch, and space environments on the ab-
sorptivity/emissivity and stability characteristics of various materi-
als/coatings commonly utilized for passive thermal control. This
data will be useful in obtaining correlation of space-environment
simulation experiments.
Design Concept.
The experiment objective will be satisfied by subjecting three dif-
ferent coating materials to a series of four exposure conditions
on the exterior of the SA-206 IU during pre-launch, launch, and
orbital operations. Each of the specimen materials will be mounted
on a temperature sensor and will be thermally isolated from sur-
rounding structure. Temperature measurements obtained from the
sensors during orbit will allow determination of thermal degrada-
tion characteristics. The degradation environments to which the
specimen thermal control coating materials will be exposed are
categorized as follows:
Pre-Launch Environment. This includes effects of moisture absorp-tion and dust accumulation immediately prior to launch.
Launch Environment. This includes heat and erosion effects during
launch ascent as well as effects of retrorocket firing and spacecraft
tower jettison.
Space Environment. One set of specimen coatings will be exposed
to all environments. A second set will be exposed to S-IB/interstage
retrorocket firing, Launch Escape System (LES) jettison and
space environment. A third set will be exposed to LES jettison
and space environment, while the fourth series will be exposed
to space environment only. The test specimens will be located at
two positions on the IU to provide two degrees of exposure to
retrorocket firing. Aerodynamic fairings will be provided to assure
a laminar flow stream for the test specimens during launch ascent.
Equipment Description and Function.
The principal elements of this experiment consist of two sensor
panels, each containing 12 thermal sensors arranged in four rows
of three. See figure 6-29. Three different thermal control coating
specimens are mounted on the sensors in each row, with each
column containing the same specimen material. Three of the four
sets of test specimens on each panel are protected with covers
Section VI Instrument Unit
attached by armament thrusters. The remaining set is protected
by a bolt-on cover which will be removed prior to launch. In order
to provide a worst-case baseline for data correlation, one specimen
on each panel is covered with a black, totally absorbing paint.
See figure 6-30.
The two sensor panels will be externally mounted on the IU with
one panel directly in line with the center of one of the S-IB/inter-
stage retrorockets for maximum exposure to heat and erosion
effects. The second panel is located approximately 51 ° around the
IU from the same retrorocket centerline for peripheral exposure.
Aerodynamic fairings are provided at each end of the sensor panels.
The experiment will utilize the power, functional command, attitudecontrol and data handling capabilities of the IU.
Operation.
Crew participation is not required for any phase of this experiment.
Although the IU telemetry system is operating, the experiment
is passive during launch ascent and orbital injection phases of the
flight except when the IU digital computer commands the expulsion
of the specimen covers by firing the armament thrusters. The
experiment functions for the duration of the S-IVB/IU lifetime
(9 to 12 hours). Experiment data is obtained after the CSM sepa-rates from the IU. The S-IVB/IU is rolled about the X axis so
that at local noon the sun-line passes through a point on the IU
midway between the two sensor panels. The X axis is maintained
along the flight vector, perpendicular to local vertical. This attitude
is maintained during data acquisition.
S150, GALACTIC X-RAY MAPPING.
The objectives of experiment S150 are to survey a portion of the
sky for X-rays in the 200 to 12,000 electron-voh energy range and
determine the intensity, location, and spectrum of each X-ray
source. In addition, the experiment will determine if there is a
continuous background of X-rays or merely discrete sources. The
results of the experiment will help answer questions concerning
emission mechanisms in X-ray sources, the distance to the sources,and the nature of the interstellar medium.
This experiment was conceived to bridge the data gap betweensounding rockets and long-life satellites in the field of X-ray as-
tronomy.
The experiment will be initiated at the completion of the SA-208
vehicle primary mission and after separation of the command and
service module. The IU/S-IVB will be oriented with the experiment
assembly (Location 24) nearest the earth. Then the experiment
sensor will be deployed (figure 6-31) so that the base plane of
the experiment sensor casting makes an angle of 62-1/2 ° with
respect to the IU/S-IVB longitudinal axis, swinging away from
its stowed position on a specially designed, hinged bracket assembly.
An onboard storage and regulation system will be used to supply
PI0 gas (10% methane, 90% argon) to the X-ray counters in the
experiment assembly. An Auxiliary Storage and Playback (ASAP)
package, consisting of five components mounted at Location
23, will record and play back data once per orbit when the vehicle
is in communications contact with a ground station. Radioactive
sources will make possible calibration of the instruments in flight.
Star sensors will help determine the pointing direction of the pro-
portional counters.
IU/S-IVB roll maneuvers will be used to allow viewing a different
portion of the sky on each orbit. Guidance requirements dictate
that the IU remain active throughout the experiment. Hence, the
maximum duration of the experiment is determined by the duration
of IU expendables. Other limiting factors might be the amount
of PI0 gas available or the quantity of thruster fuel, but it is
anticipated that the experiment will operate for at least three orbits
and possibly for five orbits.
6-29
Section VI Instrument Unit
M415 SENSORPANELS
J_ TEMPERATURE
/ SENSORLEAOS/ _" COVER RELEASE
/ / T._S_ER(6PLACES)LAUNCH / /
FAI RING
J I =,_,_, _<_,0,ONE,CO,, / _R_.__J "_-.-,LOW-O FF COVERS
FAtRING J (12 PLACES)
POS IV
POS III
SENSOR PANEL 2
_-- sENSOR PANEL 1
POS I
C-H 20035
Figure 6-296-30
M41 SENSOR MATERIALS
LAUNCH VECTOR
COVER _ I_
,...®.¢.®.] ..©¢®.]"©®®l '®©®Jl
i[¢[©[®.l I©®¢1
PANEL 1 PANEL 2
COVER REMOVAL TIME
1 COVERS REMOVED ON PAD' JUST PRIOR TO LAUNCH
2 COVERS REMOVED BY IU SEQUENCEDSQUIBS JUST PRIOR TO RETROFIRE
COVERS REMOVED BY IU SEQUENCED3 SQUIBS JUST PRIOR TO LES TOWER JETTISON
4 COVERS REMOVED BY IU SEQUENCED SQUIBSAFTER CSM SEPARATION
SPECIMAN MATERIAL
S-13G, ZINC-OXIDE PIGMENT IN METHYL/SILICONEA. BINDER ROUGH SURFACE
B. Z-93, ZINC-OXIDE PIGMENT IN SODIUM SILICATEBINDER MEDIUM ROUGH SURFACE
HXW, MSFC COMPOSITE OF SYNTHETIC MICA, ALSOC. CONTAINS POTASSIUM SILICATE AND ZI NC-OXIDE
D. THiS SPECIMEN BC BLACK CONTROL,CAT-A-LAC BLACK MED ROUGH SURFACE
C.H 20036
Figure 6-30
$150 EXPERIMENT SENSOR DEPLOYED
PROTECTIVE COVER
C-H 20037
Figure 6-31
MAJOR DIFFERENCES.
The following paragraphs summarize the differences in con-
figuration and function between the Skylab Saturn IB Instru-
ment Units and the Instrument Unit flown on the AS-205 mis-sion.
Section Vl Instrument Unit
STRUCTURAL SUBSYSTEM.
Utilize Saturn V ground half umbilical housings to launchSaturn IB vehicles
This configuration Will yield the more reliable four-ball-lock
mechanism. Implementation of this change allows usage of
the Saturn V type housing to launch IB vehicles scheduled
to be launched from Saturn V launch complex 3g facilities.
ENVIRONMENTAL CONTROL SUBSYSTEM.
a. Added a 20 micron filter in the inlet port of the gas.bearing
regulator
To preclude contamination of the gas-bearing regulator, a
20-micron filter is required.
b. Methanol/water coolant changed to Oronfte FIo-Cool 100
Oronite FIo-Cool 100 is less susceptible to corrosive galvaniccurrents and will reduce the alkaline content which reacts with
the cooling system as compared to Methanol/Water.
c. Enlarged volume coolant accumulator
To allow sufficient volume to accommodate Oronite leakage
replacement and increased volume due to its expansion, an
enlarged volume coolant accumulator is required.
d. Installation of hydraulic snubber assembly (installed be-
tween ECS coolant pump outlet and pressure transducer).
Tests have demonstrated that a hydraulic snubber located
between the ECS coolant pump outlet and pressure transducer
will prevent transducer from responding to dynamic pressure
fluctuations from the ECS coolant pump. To extend the fife
of the pressure transducer by reducing its cycling, a snubber
is required.
e. ON,storage spheres changed from two 819 cm 3 (50 in3)
to one 2704 cm 3 (165 ins).
Increased ON2 volume for longer duration mission. To meet
the 6.8-hour mission duration with a safety factor of 1.5, addi-
tional ON2 pressurant is required.
f. Removal of modulating flow control valve, electronic con-
troller assembly, and associated hardware.
Since the deletion of usage of the modulating flow control
valve during ground operations and because it is inoperative
during flight, the requirement for this hardware is deleted.
g. Delete usage of hazardous gas detection system. With the
addition of sample ports to the S-IVB stage for Saturn IB
vehicles, the requirement for the IU hazardous gas systemis deleted.
h. The IU redundant pump system and related hardware have
been Incorporated into the design for S-IU-206 through S-IU.
212 and S-IU-502 through S-lU-515.
To efiminate the coolant pump as a single point failure, an
additional coolant pump is required.
GUIDANCE SUBSYSTEM.
There were no significant changes made directly to the guid-
ance components; however, the following S-IU-206 changes
were made that indirectly affect the guidance subsystem.
a. The IU was modified to allow backup guidance commands
from the spacecraft platform and computer during first stage
burn and manual guidance commands from the spacecraft
during second stage burn. In addition, the change imposed
6-31
Section Vl Instrument Unit
the requirement that an IU guidance failure must occur before
enabling manual control during boost.
b. The IU was modified to provide dual redundant power inputs
to components of the ST.124M.3.
FLIGHT CONTROL SUBSYSTEM.
a. The coolant lines on the S-IU-206 FCC are modified by
the addition of an innersleeve.
The innersleeve provides a redundant line in case of a coolant
line flare cracking. During QAST of the Saturn V FCC, the
flared tubing connecting the front panel coolant ports to the
FCC spider cracked in the flared area. Failure analysis revealed
that the cracked flares were primarily the result of slow cycle
fatigue compounded by high flare stress due to tube misalign-
ment, stress riser due to out-of-tolerance flare radius, or flexingof the flare radius due to normal relative motion between the
spider and front panel.
b. Flight control computer spatial filters were added, and also
the FCC filters, used in boost flight, were changed to enhance
the vehicle dynamic performance.
Analysis of previous flight data, specifically AS-204, indicated
that during the S-IVB burn phase a high level ac rate signal
was present on the FCC input. The ac signal frequency was
approximately 18 Hz and reached an ampfitude of 1.75 degrees
per second. The overall effect of this condition resulted in
a decreased attitude error sensitivity, thus causing the APS
engines to fire only with attitude error signals significantly
larger than the deadband of the spatial amplifiers.
c. Implement Saturn V backup guidance scheme into SaturnIB FCC.
This change redesigned the FCC inputs so that spacecraft
control could be attained during the S-IB burn phase.
The S-IU-205 FCC did not contain the capability for control
from the spacecraft during S-IB burn. Additionally, if spacecraft
control were energized during S-IVB burn, the error signals
were attenuated and limited by the limiter circuits within the
FCC.
ECR AA OE-212 established the requirement for manual backup
guidance during all modes of flight. In the AS-205 configu-
ration, no backup guidance existed during the S-IB burn mode.This meant an abort condition existed, if the normal IU guid-
ance system had a failure in a non-redundant circuit. The
requirement of a gain of 20 during spacecraft control on the
roll input DC amplifier is required to make the FCC compatible
with spacecraft scale factor for the roll channel
d. A modification to the FCC inverter detection, to prevent
switching during power transfer, was implemented.
The AS-205 inverter detector was sensitive to negative voltage
transients on the battery buses feeding the FCC. During the
AS-204 and AS-205 countdowns, the inverter detector tripped
during the IU power transfer test at T-30 minutes. The tripwas caused by the voltage of the battery peaking while un-
loaded.
The launch mission rules require that the detector be reset
at launch. If the detector trips during the last power transfer,
a recycle to T-20 minutes is required. If this occurred duringthe SL-2 countdown, the launch would be slipped to the fol-
lowing day due to the short launch window for rendezvous.
Therefore, the detector was redesigned to make it insensitive
to voltage transients on the input voltage.
e. Addition of mechanical support to the S-IB FCC center
tray
6-32
The Saturn IB FCC failed the QAST conducted on November
22, 1971, by IBM. The failure created a loss of 12 electrical
signals being monitored on recorders. The subsequent failure
analysis revealed five structural discrepancies within the unit.
f. Implement control gains and shaping networks for S-IB FCC
The change involves modifying the gains of these networks
for the first 100 seconds of S-IB Stage flight time. The following
table shows the "was-is" for each of the control gains:
Was Is
ao 1.55 1.85
a, 1.70 1.65
g_ 5.0 4.0
The gain change is accomplished by changing resistor values
in the networks.
This change was implemented as a result of the AS-206/SL-2
phase II control system design and analysis effort. The report
indicated the desirability of increasing the aerodynamic gain
margin in the region of maximum aerodynamic pressure (max
Q). The gain changes provide a greater aerodynamic gain
margin and thereby increase confidence in the ability of the
vehicle to accomplish the mission even with severe off-nominal
conditions such as an engine out or degradation due to para-meter variation.
FLIGHT PROGRAM.
The program has been re-written for the purpose of modulari-
zation to make a generalized flight program adaptable to mod-ifications which are mission dependent. Internal functional and
mission requirement changes are implemented through
changes in processing within the module which performs the
function or parameter changes in tables which are processed
for the purpose of computation or timing.
The LVDC Equation Defining Document for the Saturn IB Flight
Programs should be consulted for information concerningmission dependent requirements as well as general require-
ments.
ELECTRICAL SUBSYSTEM.
a. S.IB engine cut-off circuitry redesign.
Engine cutoff circuits redesigned to eliminate all single pointelectrical failures which could cause an inadvertent multiple
engine cutoff. Also improved reliability and range safety cutoff.
b. Implement redundancy to enhance probability of comple-
ting the prime mission in case of 6D10 or 6D30 battery failure.
The 6D10 battery provided the only source of power for theswitch selector and certain control functions. In addition,
switch point power for the FCC was supplied only by + 6D31bus. In the event of battery failure, the following unsatisfactory
conditions could have developed:
Spacecraft control of Saturn could not be attained
Switch selector commands would not be issued
S-IB outboard engines cutoff, S-IB level sensors dry, and
S-IVB engine out "A" and "B" discretes and interruptsinto the L VDA would not be issued
The flight control computer switch points were dependent
on + 6D31 power and loss of the + 6D30 battery would
result in incorrect FCC channel gains
c. Implement redundant power for ST-124M-3 stabilized plat-
form system.
In order to meet the redundant power requirements to the
platform, the method of paralleling the output of the battery
buses where required (+6DI0 and +6D30) with diodes at
the input to the various platform subsystems was chosen
rather than attempting to parallel the total output of both
batteries with a single set of diodes.
This change provides redundant power to the 56-volt power
supply with a separate set of bus isolation diodes, redundant
power to the platform AC power supply with another separateset of isolation diodes, and redundant power to the platform
electronic assembly with the existing + 6D61 bus (redundant
power bus).
Battery + 6D10 was the single source for the ST-124M-3 prime
power.
d. S-IVB passivation will fly in the S.IB vehicles for the firsttime onboard S-IU-206.
This change will allow the dumping of S-IVB propellants after
spacecraft/launch vehicle separation, thereby sating the spent
stage. Prior to this change, the S-IVB cutoff signals from the
IU locked-out any possibility of dumping the propellants.
e. Addition of telemetry measurement K91-602 allows the
monitoring of the redundant EDS cutoff command.
This change converts from an unmanned configuration to a
manned configuration.
f. Implement redundant paths for IU/ESE launch critical func-
tions.
The IU has critical functions through the IU/ESE interface
connectors which if lost due to an open circuit would cause
an engine shutdown after ignition. Other critical functions
through the IU/ESE interface connectors if open could prevent
a desired engine shutdown after ignition. Due to the vibration
from engine ignition until liftoff and the possibility of faultyESE connectors or of contaminants in the vehicle umbilical
connectors, open circuits could occur.
g. Implement redundancy for TB-2 and TB-3 initiate.
The primary indications to the LVDA/LVDC for initiation of
time bases 2 and 3 are interstage functions and do not have
hardware backups. This change provides redundancy for these
critical interstage functions.
h. Deletion of fourth battery (6D20).
Reconfiguration and redesign resulted in no requirement.
INSTRUMENTATION AND COMMUNICATIONS
SUBSYSTEM.
a. Implement flight control measurements rechannelization.
There are additional instrumentation and communications
subsystem design requirements resulting from the S-IU-206
Section VI Instrument Unit
role in supporting the SL-2 mission. New design requirements
are as follows:
Flight control measurements shall appear on the sametelemetry channels on S-IU-206, S-IU-207, S-IU-208, and
S-IU-513. This required rechannelization of ten measure-
ments on S-IU-206.
b. PCM telemetry design changes (register switch, PSR and
digital gate).
Existing PCM telemetry system is highly susceptible to exces-sive IU and/or vehicle electrical noise environments. Redesign
of the Reg. Sw. PSR and digital gate will provide a more nearly
optimum noise free PCM telemetry system.
c. Implement thermal control coating experiment (M-415).
The IU I&C subsystem shall be modified to provide signal
conditioning and transmission of experiment data. Signal con-
ditioning equipment shall include differential amplifiers and
divider networks for a total of 26 resistance thermistors.
Timing shaft be provided with all experiment telemetry datato allow correlation with vehicle orbital data and to allow plot-
ting of data as a function of time.
OVERALL RELIABILITY/PRODUCT
IMPROVEMENT.
Modify/change hardware and circuitry to make SA-206 thru
SA.212 man-rated. (S-IB phase-up).
1. Rework included addition of Q-balls, addition of EDS payload
power (+ 6D91, + 6D92 and + 6D93), addition of EDS redun-
dant power, change in EDS rate switch circuits, etc.
2. Added battery temperature measurements
3. S-IB Phase-Up, EDS and LV/CSM Electrical Interface
4. Redesign of IU command decoder for solder joint stress
relief
5. Implement Changes to RDM Power Supply Assembly
6. Adjust pressure switch setpoint
7. Correct undesirable turnoff of coolant pump number 1
8. Provide line to shut-off coolant pum_ number 2
9. Desensitize coolant pump pressure switch
10. Implement redesigned water accumulator diaphragm
11. Implement redesigned gas-bearing heat exchanger
12. Improved O-rings for ECS coolant loop
13. Lightning detection devices and associated temporary
cabfing has been added for monitoring during pre-launchcheckout. All cabling will be removed just prior to launch.
6-33
TABLE OF CONTENTS cells, four of which are equipped with systems to simulate interface
Introduction .................................................................... 7-1
Vehicle Assembly Building ............................................ 7-1Launch Control Center .................................................. 7-1
Mobile Launcher ............................................................ 7-3
Launch Pad .................................................................... 7-10
Mobile Service Structure .............................................. 7-14
Crawler-Transporter ........................................................ 7-14
Converter-Compressor Facility ...................................... 7-14
Ordnance Storage Area ................................................ 7-16
Vehicle Assembly and Checkout .................................. 7-16
Hold and Recycle Criteria ............................................... 7-22Launch Interlocks ........................................................... 7-22
INTRODUCTION.
Launch Complex 39 (LC-39), Kennedy Space Center, Florida,
which originally provided all the facilities necessary to the assembly,
checkout, and launch of the Apollo/Saturn space vehicle has been
modified to accommodate the Skylab Saturn IB and Saturn V space
vehicles. The vehicle assembly building (VAB) provides a controlledenvironment in which the vehicle is assembled and checked out
on a mobile launcher (ML). The space vehicle and the launch
structure are then moved as a unit by the crawler-transporter to
the launch site, where vehicle launch is accomplished after propel-
lant loading and final checkout. The major elements of the launch
complex shown in figure 7-1, are the vehicle assembly building
(VAB), the launch control center (LCC), the mobile launcher (ML),
the crawler-transporter (C-T), the crawlerway, the mobile service
structure (MSS), and the launch pad.
VEHICLE ASSEMBLY BUILDING.
The VAB is located adjacent to Kennedy Parkway, about five miles
north of the KSC industrial area. Its purpose is to provide a
protected environment for receipt and checkout of the propulsion
stages and instrument unit (IU), erection of the vehicle stages and
spacecraft in a vertical position on the ML, and integrated checkout
of the assembled space vehicle.
The VAB, as shown in figure 7-2 is a totally enclosed structure
covering eight acres of ground. It is a structural steel building
approximately 525 feet high, 518 feet wide, and 716 feet long.
The siding is insulated aluminum except where translucent fibel-
glass sandwich panels are used in part of the north and southwalls.
The principal operational elements of the VAB are the low bay
area and high bay area. A 92-foot wide transfer aisle extends
through the length of the VAB and divides the low and high bay
areas into equal segments (figure 7-3).
LOW BAY AREA.
The low bay area provides the facilities for receiving, uncrating,checkout, and preparation of the S-IVB stage, and the IU (and
S-ll stage for Saturn V). The low bay area, located in the southernsection of the VAB, is approximately 210 feet high, 442 feet wide,
and 274 feet long. There are eight stage preparation and checkout
operations between the stages and the IU.
Work platforms, made up of fixed and folded sections, fit about
the various sections as required. The platforms are bolted, to permit
vertical repositioning, to the low bay structure. Access from fixed
floor levels to the work platforms is provided by stairs.
HIGH BAY AREA.
The high bay area provides the facilities for erection and checkout
of stages and spacecraft, and integrated checkout of the assembled
space vehicle. The high bay area which is located in the northern
section of the building, is approximately 525 feet high, 518 feet
wide, and 442 feet long. It contains four checkout bays, each capable
of accommodating a fully assembled space vehicle.
Access to the vehicle at Various levels is provided from air condi-
tioned work platforms that extend from either side of the bay to
completely surround the launch vehicle. Each platform is composed
of two biparting sections which can be positioned in the vertical
plane. The floor and roof of each section conform to and surround
the vehicle. Hollow seals on the floor and roof of the section provide
an environmental seal between the vehicle and the platform.
Each pair of opposite checkout bays is served by a 250-ton bridge
crane with a hook height of 462 feet. The wall framing between
the bays and the transfer aisle is open above the 190-foot elevation
to permit movement of components from the transfer aisle to their
assembly position in the checkout bay.
The high bay doors provide an inverted T-shaped opening 456
feet in height. The lower portion of the opening is closed by doorswhich move horizontally on tracks. The upper portion of the open-
ing is dosed by seven vertically moving doors.
UTILITY ANNEX.
The uiility annex, located on the west side of the VAB, supports
the VAB, LCC and other facilities in the VAB area. It provides
air conditioning, hot water, compressed air, water for fire protection,
and emergency electrical power.
HELIUM/NITROGEN STORAGE-VAB AREA.
The gas storage facility at the VAB provides high pressure gaseous
helium and nitrogen. It is located east of the VAB and south of
the crawlerway. The roof deck of the building is removable to
permit installation and removal of pressure vessels through the
roof. This facility is serviced from the converter-compressor facility
by a 6,000 psig gaseous helium line and a 6,000 psig gaseous
nitrogen line.
LAUNCH CONTROL CENTER.
The LCC (figure 7-4) serves as the focal point tot overall direction,
control, and surveillance of space vehicle checkout and launch.
The LCC is located adjacent to the VAB and at a sufficient distance
from the launch pad (three miles) to permit the safe viewing of
liftoff without requiring site hardening. An enclosed personnel and
cabling bridge connects the VAB and LCC at the third floor level.
The LCC is a four-story structure approximately 380 by 180 feet.
The ground floor is devoted to service and support functions such
7-1
Section VII Ground Support Interface
LAUNCH COMPLEX 39
LAUNCH AREA B
(SATURN I B)
CH AREA A
(SATURN V)
C RAWLERWAY
MOBILE SERVICE
STRUCTURE PARK
ORDNANCE
CONVE_ER/
COMPRESSOR
MOBILE LAUNCHER
REFURBISH
VEHICLE ASSEMBLY
BUILDING
STORAGE
TURNING BASIN
LAUNCH CONTROL CENTER
BARGE CANAL
7-2 Figure 7-1
as cafeteria, offices, shops, laboratories, the communications control
room; and the complex control center. The second floor houses
telemetry, RF and tracking equipment, in addition to instrumenta-tion and data reduction facilities.
The third floor is divided into four separate but similar control
areas, each containing a firing room, computer room, mission
control room, test conductor platform area, visitor gallery, offices
and frame rooms. Three of the four firing rooms contain control,
monitoring and display equipment for automatic vehicle checkoutand launch.
Direct viewing of the firing rooms and the launch area is possible
from the mezzanine level through specially designed, laminated,
and tinted glass windows. Electrically controlled sun louvers are
positioned outside the windows.
The display rooms, offices, launch information exchange facility
(LIEF) rooms, and mechanical equipment are located on the fourthfloor.
The electronic equipment areas of the second and third floors haveraised false floors to accommodate interconnecting cables and air
conditioning ducts.
The power demands in this area are large and are supplie d by
two separate systems, industrial and instrumentation. The industrial
power system supplies electric power for lighting, general use recep-
tacles, and industrial units such as air conditioning, elevators, pumps
and compressors. The instrumentation power system supplies powerto the electronic equipment, computers, and related checkout
equipment. This division between power systems is designed to
protect the instrumentation power system from the adverse effects
of switching transients, large cycling loads, and intermittent motor
starting loads. Communication and signal cable provisions have
been incorporated into the design of the facility. Cable troughsextend from the LCC via the enclosed bridge to each ML location
in the VAB high bay area. The LCC is also connected by buriedcableways to the ML refurbishing area and to the pad terminal
connection room (PTCR) at the launch pad. Antennas on the roof
provide an RF link to the launch pads and other facilities at KSC.
Section VII Ground Support Interface
MOBILE LAUNCHER.
The mobile launcher (figure 7-5) is a transportable steel structure
which, with the crawler-transporter, provides the capability to move
the erected vehicle to the launch pad. The ML is divided intotwo functional areas, the launcher base and the umbilical tower.
The launcher base is the platform on which a Saturn V vehicle
is assembled in the vertical position, transported to a launch site,
and launched. The umbilical tower, permanently erected on the
base, is the means of ready access to all important levels of the
vehicle during the assembly, checkout, and servicing periods prior
to launch. The equipment used in the servicing, checkout, and
launch is installed throughout both the base and tower sections
of the ML. The intricate vehicle-to-ground interfaces are established
and debugged in the convenient and protected environment of theVAB, and moved undisturbed aboard the ML to the pad.
Mobile launcher 1 is used for Skylab Saturn IB space vehicles,
and has been modified by (1) addition of a pedestal, to maintain
the spacecraft and upper stages' interfaces with the umbilical tower
service arms; and (2) removal or inactivation of S-II and S-IC
stage GSE and addition of S-IB stage GSE.
LAUNCHER BASE.
The launcher base is a two story steel structure 25 feet high, 160
feet long, and 135 feet wide. Each of the three levels provides
approximately 12,000 square feet of floor space. The upper deckis designated level O. Level A, the upper of the two internal levels,
contains 21 compartments and level B has 22 compartments.
A new structural pedestal is installed on the ML deck that raises
the Saturn IB launcher platform to the 127 foot level. The basic
tower is an open truss welded steel pipe structure. An access bridge
is provided from the pedestal deck level to the ML tower. Firing
accessories for the S-IB stage were remowed from LC-34/37B and
are installed on the pedestal deck. There is an opening through
the pedestal deck centered over the opening in the ML base for
VEHICLE ASSEMBLY BUILDING
VEHICLE
ASSEMBLY
BUILDING
LAUNCH
CONTROL
CENTER
Figure 7-2
C-H 20039
7-3
Section VII Ground Support Interface
first stage exhaust. A work platform is provided for the pedestal
opening for prelaunch engine servicing. Pneumatics, propellants,
water, and other services are routed to the pedestal base, with valve
panels installed in an enclosed equipment level below the pedestaldeck.
Access to the launcher base interior is provided by personnel/
equipment access doors opening into levels A and B and equipmentaccess hatches located on levels O and A.
The base has provisions for attachment to the crawler-transporter,
six launcher-to-ground mount mechanisms, and four extensiblesupport columns.
All electrical/mechanical interfaces between vehicle systems and
the VAB or the launch site are located through or adjacent to
the base structure. A number of permanent pedestals at the launch
site provide support for the interface plates and servicing lines.
The base houses such _tems as the computer systems test sets,
propellant loading equipment, hydraulic test sets, propellant and
pneumatic lines, air conditioning and ventilating systems, electrical
power systems, and water systems. Shock-mounted floors and spring
supports are provided so that critical equipment receives less than
-----0.5 G mechanically-induced vibrations. Electronic compartments
within the ML base are provided with acoustical isolation to reduce
the overall rocket engine noise level.
The air conditioning and ventilating system for the base provides
environmental protection for the equipment during operations and
standby. One packaged air conditioner provides minimal environ-
mental conditioning and humidity control during transit. Fueling
operations at the launch area require that the compartments within
the structure be pressurized to a pressure of three inches of water
above atmospheric pressure and that the air supply originate froma remote area free from contamination.
The primary electrical power supplied to the ML is divided intofour separate services: instrumentation, industrial, in-transit and
emergency. Instrumentation and industrial power systems are sepa-
rate and distinct. During transit, power from the crawler-transporter
is used for the water/glycol systems, computer air conditioning,
threshold lighting, and obstruction lights. Emergency power for
VAB INTERIOR LAYOUT
k LOW BAY AREA m
TOWER AO
AREA L AREA K i l
I o
_ SL'SrAGEARE2---
TRANSFER AISLE
s-;,STAGEAREA--;-- n--T-n
[" T=-7' I__J !
BAY 2
BAY 1
ML-1
r_r T TI-,
S-IB HOLDDOWN
PAOS
n
13I ILL.IJ
i_ SKYLA8 2
HIGH BAY AREA
STORAGE
APOLLO/SATURN V
7-4
AREA N AREA M
TOWER D
Figure 7-3
r
!__TOWER F
C-H 20040
the ML is supplied by a diesel-driven generator located in the
ground facilities. It is used for obstruction lights, emergency lighting,
and for one tower elevator. Water is supplied to the ML at the
VAB and at the pad for fire, industrial and domestic purposes
and at the refurbishment area for domestic purposes.
FIRING ACCESSORIES.
The ML firing accessories (figure 7-6) described in the following
paragraphs were removed from LC-34/LC-37B and installed on
the pedestal deck.
Holddown Arms.
The holddown arms system (figure 7-6) provides support for the
vehicle on the launch pedestal and restrains the vehicle from flight
until verification of full thrust at launch. The holddown arms _y=:,_m
consists basically of eight holddown arms, .'_=-'.Ahoiadown release
mechanisms, pneumatic separators, and the necessary pneumatic
control panel. The pneumatic system is redundant from the inletsupply line on the control panel through to the pneumatic separa-
tors. The supply panel, once charged with helium, has a reservoir
with the capacity to effect release of the vehicle upon command,
even with complete failure of the inlet supply system. The solenoids
release gaseous helium at 750 psig which pressurizes the separators
through redundant "tuned length" tubing. The tubing is tuned
length; that is, the distance and the volume are identical from
the solenoid valves to each and all of the pneumatic separators.
The criteria here is not total elapsed time but the total time dif-ference between first arm release to last arm release, and the use
of helium minimized this time lag. The time from release command
to the slowest arm retraction is approximately 140 msec. The activa-
tion pressure required to release the arms is 100 psig. An explosivebolt assembly is added to each of the eight holddown arm pneu-
matic separators. This backup system provides for the activation
of the explosive release device if first movement of any or all armsis not received within a specific time period (approximately 190
msec) from the initiation of the release command. The basic hold-
down arm system has verified its reliability during previous
launches. The backup system consists of an explosive bolt assembly
that increases the reliability factor.
Section VII Ground Support interface
Boattail Conditioning Lines.
The bottail conditioning and water quench system installation
provides the final link connecting the vehicle to the boattail condi-
tioning and water quench system. The water quench system
transfers water at a rate of 8000 gpm at 125-psig pressure to the
vehicle boattail area when necessary for combating fires prior tovehicle liftoff. In addition, the installation is used in conjunction
with the environmental control system and deluge purge system
to provide a controlled atmosphere in the boattail area for personnel
safety and vehicle protection during various periods while the
vehicle is on the launch pedestal. The controlled atmosphere is
provided by two purges: air purge and nitrogen (GNu) deluge purge.The installation automatically disconnects at liftoff after approxi-
mately 2-1/4 in. of vehicle travel. The boattail conditioning and
water quench system installation is essentially four installations,
each in general consisting of a support bracket, disconnect coupling,
flexible hose assembly, valve assembly, elbow assembly, and the
hardware required for installation. The four installations are in-
stalled at the four principal positions and are similar. The flexible
hose assembly is covered with a heat-and-blast-protective coating.
Fuel Fill Mast.
The fuel fill mast system is used to provide a means for connectingthe fuel transfer line to the launch vehicle and to provide a fast,
fail-safe method for disconnecting the line at liftoff. The mast
supplies fuel to the vehicle at a pressure of 50 psig and a flowrate
of 1500 gpm. The pivotal feature of the fuel fill mast providesfor mast erection and retraction, thereby providing adequate liftoffclearance for the Vehicle. Mast retraction at vehicle liftoff is an
automatic function tied to the firing circuits for proper sequencing;
however, manual control for erecting and testing can be performed
through use of a pneumatic valve box mounted on the cross member
of each mast support stand. The fuel fill mast consists of a
retractable coupling assembly, two cylinder assemblies, support
bracket assembly, retracting assembly, mast arrestor mounting
assembly, valve box lox assembly, support stand, upper and lower
pipe weldments, and a hose assembly that provides a flow path
for the fuel during propellant tanking operation. The fuel fill mast
Figure 7-4
C.H 20041
7-5
Section VII Ground Support Interface
MOBILELAUNCHER
UMBILICAL TOWER_
, _ HAMMERHEAD CRANE
PEDESTAL
LAUNCHER BAS E---'_
CH-20042
7-6 Figure 7-5
incorporates provisions for vertical, lateral, and retraction adjust-
ments to ensure correct alignment and operation.
Lox Fill Mast.
The lox fill mast system provides a means of connecting the
lox transfer line to the launch vehicle and provides a fast, fail-safe
method for disconnecting the line at vehicle liftoff. Disconnectionof the lox transfer line at liftoff is a function of vehicle motion,
as the vehicle simply lifts away from the mast assembly. Mastretraction after vehicle liftoff is a remotely controlled function of
the pneumatic circuits. These circuits are controlled by solenoidvalves located in the launcher valve box. The lox fill mast system
functional requirements are to supply liquid oxygen to the vehicle
through a 6-in. coupling at a pressure of 90 psig with a flow rate
Section VII Ground Support Interface
of 1250 gpm, with inflight disconnect capability. The lox fill mastconsists of a retractable coupling assembly, two cylinder assemblies,
support bracket assembly, retracting assembly, mast arrestor,
mounting assembly, valve box assembly, support stand, upper and
lower pipe weldments, hose assembly and pneumatic lines. The
retractable coupling assembly, upper and lower pipe weldments,
and hose assembly provide a flow path for the lox during propellant
tanking operations. The lox fill mast incorporates provisions for
vertical, lateral, and retraction adjustments to ensure correct align-
ment and operation.
Short Cable Masts.
The short cable masts provide a connecting link, structural sup-
port, and disconnecting capability for electrical cable and pneumatic
POS. III _ .
\
MAST
SHORT CABLE MAST
POS. II
/POS. IV
MAST
SERVICE ARM
CONTROL
SWITCH
RT CABLE MAST
POS. IV
/iI
ERVICE ARM /
CONTROL SWITCH ¢
ARM
/ (8 pLACES)
BOATTAIL
CONDITIONING
(4 PLACES)
CH.2O043
Figure 7-6
7-7
Section VII Ground Support Interface
S-IB FORWARD (INFLIGHT). PROVIDES PNEU-
MATIC, ELECTRICAL, AND AIR-CONDITIONING
INTERFACES. UMBILICAL WITHDRAWAL BY PNEU-
MATIC DISCONNECT IN CONJUNCTION WITH
PNEUMATICALLY DRIVEN BLOCK AND TACKLE/
LANYARD DEVICE. SECONDARY MECHANICAL
SYSTEM. RETRACT TIME IS 3.9 SECONDS.
S-IVBAFT (INFLIGHT). PROVIDES LH 2AND
LOX TRANSFER, ELECTRICAL, PNEUMATIC,
AND AIR-CONDITIONING INTERFACES.
UMBILICAL WITHDRAWAL SYSTEMS SAME
AS S-IVB FORWARD. ALSO EQUIPPED WITH
LINE HANDLING DEVICE. RETRACT TIME IS
6. I SECONDS (MAX).
S-IVB FORWARD/IU UMBILICAL (INFLIGHT).
FOR S-IVB STAGE, PROVIDES FUEL TANK VENT,
ELECTRICAL, PNEUMATIC, AIR-CONDITIONING,
AND PREFLIGHT CONDITIONING iNTERFACES.
FOR IU, PROVIDES PNEUMATIC, ELECTRICAL, AND
AIR-CONDITIONING INTERFACES. UMBILICAL
WITHDRAWAL BY PNEUMATIC DISCONNECT IN
CONJUNCTION WITH PNEUMATIC/HYDRAULIC
REDUNDANT DUAL CYLINDER SYSTEM. SECOND-
ARY MECHANICAL SYSTEM. ARM ALSO EQUIPPED
WITH LINE HANDLING DEVICE TO PROTECT LINES
DURING WITHDRAWAL. RETRACT TIME IS 6. I SEC-
ONDS (MAX).
SERVICE MODULE (INFLIGHT). PROVIDES AIR-
CONDITIONING, VENT LINE, COOLANT,
ELECTRICAL, AND PNEUMATIC INTERFACES.
UMBILICAL WITHDRAWAL BY PNEUMATIC/
MECHANICAL LANYARD SYSTEM WITH SECOND-
ARY MECHANICAL SYSTEM. RETRACT TIME IS
6. I SECONDS (MAX).
COMMAND MODULE ACCESS ARM (PREFLIGHT).
PROVIDES ACCESS TO SPACECRAFT THROUGH
ENVIRONMENTAL CHAMBER. ARM MAY BE
RETRACTED OR EXTENDED FROM LCC. RETRACTED
TO 12 ° PARK POSITION DURING PERIOD T-43 TO
T-5 MINUTES. EXTEND TIME IS 12 SECONDS FROM
THIS POSITION.
CH-20044
7-8 Figure 7-7
service lines required for checkout and operation of Saturn IB
vehicle engine components prior to and during the process oflaunch. After the vehicle has lifted approximately four inches off
the launch pad, the masts disconnect all cables and lines, and retract
from the vehicle. Each short cable mast consists of a support
platform, a mast weldment, two kickoff cylinders, a retract cylinder,
a quick-release housing, a latch-back mechanism, electrical cables,
pneumatic service lines, and supporting hardware. The double-ac-
tion, pneumatic retract cylinder furnishes the primary force (or
retracting the short cable mast during launch. A 750-psig GN z
pressure is applied to the top of the cylinder while an opposing
constantly venting 50-psig GN 2 pressure is applied at the bottom
of the cylinder and acts as a cushioning pressure for the mast at
the end of retraction. Electrical cables are routed through an 8-in.
flexible shield to the quick-release housing. These cables connect
electrical circuitry to the S-IB stage engine compartment from the
beginning of checkout until the mast is disconnected after vehicle
liftoff. The pneumatic lines are mounted along the left and rightsides of the mast weldment. These lines supply pneumatic pressure
to the S-IB stage for standby, checkout, and launch.
UMBILICAL TOWER.
The umbilical tower is an open steel structure 380 feet high which
provides the support for umbilical service arms, access arm, work
and access platforms, distribution equipment for the propellant,
pneumatic, electrical and instrumentation subsystems, and other
ground support equipment. The distance from the vertical centerlineof the tower to the vertical centerline of the vehicle is approximately80 feet. The distance from the nearest vehicle column of the tower
to the vertical centerline of the vehicle is approximately 60 feet.
Two high speed elevators service 18 landings, from level A of thebase to the 340-foot tower level.
Section VII Ground Support Interface
The hammerhead crane is located on top of the umbilical tower.
The load capacity of the crane is 25 tons with the hook extended
up to 50 feet from the tower centerline. With the hook extendedbetween 50 and 85 feet from the tower centerline, the load capacity
is 10 tons. The hook can be raised or lowered at 30 feet per minute
for a distance of 468 feet. The trolley speed is 110 feet per minute.
The crane can rotate 360 degrees in either direction at one revo-
lution per minute. Remote control of the crane from the ground
and from each landing between levels 0 and 360 is provided by
portable plug-in type control units.
Tower modifications for Skylab Saturn IB propellant systems con-
sisted of (1) adding an RP-1 supply line from the ML base to
an RP-1 skid added at the 100 foot level, thence to the S-IB stage
fuel mast; (?) adding lox line to the S-IB stage lox mast, assignment
of the 100C gpm vacuum-jacketed lox replenish system (for Saturn
V) to accomplish both fill and replenish functions for Saturn IB,
and ina:tivation of the un-insulated 10,000 gpm lox supply linesand S_C skid used on Saturn V.
Sr_;RVICE ARMS.
The service arms provide access to the launch vehicle and support
the service lines that are required to sustain the vehicle as described
in figure 7-7. The service arms are designated as either preflight
or inflight arms. The preflight Command Module Access Arm is
retracted to a 12 deg park position during final launch preparation
and then is retracted and locked against the umbilical tower prior
to liftoff. The inflight arms retract at vehicle liftoff, after receiving
a command signal from the service arm control switches locatedin holddown arms.
The inflight service arm launch retract sequence typically consists
TYPICAL LAUNCH PAD
Figure 7-8
C-H 20045
7-9
Section VII Ground Support Interface
of the following operations: umbilical carrier release, carrier with-drawal, and arm retraction and latchback. When the vehicle rises
3/4-inch, the primary liftoff switches on the holddown arms activate
a pneumatic system which unlocks the umbilical carriers and pushes
each carrier from the vehicle. If this system fails, the secondarymechanical release mechanism will be actuated 500 milliseconds
after T-O by the service arm control switch back-up timer. Upon
carrier ejection, a double pole switch activates both the carrier
withdrawal and arm retraction systems. If this switch fails, it will
be by-passed by a signal from the backup timer. Line handlingdevices on the S-IVB forward and aft arms are also activated on
carrier ejection. Carrier withdrawal and arm retraction is accom-
plished by pneumatic and/or hydraulic systems.
Service arm modifications for Skylab Saturn IB are:SA-1A Previously an S-IC forward arm, modification consists
of making it an in-fllght arm, installing S-IB stage umbilical plates,
and adding an extension of 68" to lengthen the arm. Minor modifi-cations are made to the service arm skid. SA-IA is installed in
the former position of SA-5.
SA-6 The carrier is modified to interface with the Saturn IB
cylindrical S-IVB adapter.
SA-7 The umbilical services for the LEM are deleted.
Changes are not required of SA-8 and SA-9. The service arms
for the S-IC and S-II stages are removed (SA-I, SA-2, SA-3, SA-4,SA-5).
LAUNCH PAD.
The launch pad shown in figure 7-8 is typical of launch pads 39-A
and 39-B. The following details are applicable to both pads except
where differences are specified.
The launch pad provides a stable foundation for the ML during
launch and prelaunch operations and an interface to the ML for
ML and vehicle systems. The two pads at LC-39 are locatedapproximately three miles from the VAB area. Each launch site
is an eight-sided polygon measuring approximately 3,000 feet
across. Pad B is used for Skylab Saturn IB space vehicles.
LAUNCH PAD STRUCTURE.
The launch pad is a cellular, reinforced concrete structure with
a top elevation of 48 feet above sea level (42 feet above grade
elevation). The longitudinal axis of the pad is oriented north-south,
with the crawlerway and ramp approach from the south.
Located within the fill under the west side of thekstructure (figure7-9) is a two-story concrete building to house environmental control
and pad terminal connection equipment. On the east side of the
structure, within the fill, is a one-story coficrete building to house
the high pressure gas storage battery. On the pad surface are
elevators, staircases and interface structures to provide service to
the ML and the mobile service structure (MSS). A ramp, with
a five percent grade, provides access from the crawlerway. This
is used by the C-T to position the ML/space vehicle and the MSS
on the support pedestals. The azimuth alignment building is located
on the approach ramp in the crawlerway median strip. A flame
trench 58 feet wide by 450 feet long, bisects the pad. This trench
opens to grade at the north end. The 700,000-pound mobile
wedge-type flame deflector is mounted on rails in the trench.
An escape chute is provided to connect the ML to an underground,hardened room. This room is located in the fill area west of the
FLAME
PRESSURE
GAS STORAGE
FLOOR
APPROACH RAMP
EGRESS SYSTEM
BUILDING
-PTCR
SUBSTATION
COOLING
TOWER
C-fl 20046 i
7-10
Figure 7-9
support structure. This is used by astronauts and service crews inthe event of a malfunction during the final phase of the countdown.
PAD TERMINAL CONNECTION ROOM.
The pad terminal connection room (PTCR) (figure 7-9) providesthe terminals for communication and data link transmission con-
nections between the ML or MSS and the launch area facilities
and between the ML or MSS and the LCC. This facility also
accommodates the electronic equipment that simulates the vehicle
and the functions for checkout of the facilities during the absence
of the launcher and vehicle.
The PTCR is a tWo-story hardened structure within the fill on
the west side of the launch support structure. The launch pedestal
and the deflector area are located immediately adjacent to thisstructure. Each of the floors of this structure measures approxi-
mately 136 feet by 56 feet. Entry is made from the west side of
the launch support structure at ground level into the first floor
area. Instrumentation cabling from the PTCR extends to the ML,
MSS, high pressure gas storage battery area, lox facility, RP-t
facility, LH2 facility, and azimuth alignment building. The equip-
ment areas of this building have elevated false floors to accommo-date the instrumentation and communication cables used for inter-
connecting instrumentation racks and terminal distributors.
The air conditioning system, located on the PTCR ground floor,
provides a controlled environment for personnel and equipment.The air conditioning system is controlled remotely from the LCC
when personnel are evacuated for launch. This system provideschilled water for the air handling units located in the equipment
compartments of the ML. A hydraulic elevator serves the two floors
and the pad level.
Section VII Ground Support Interface
Industrial and instrumentation power is supplied from a nearby
substation.
ENVIRONMENT CONTROL SYSTEM.
The ECS room located in the pad fill west of the pad structure
and north of the PTCR (figure 7-9) houses the equipment which
furnishes temperature and/or humidity controlled air or nitrogen
for space vehicle cooling at the pad. The ECS room is 96 feet
wide by 112 feet long and houses air and nitrogen handling units,
liquid chillers, air compressors, a 3000-gallon water-glycol storagetank and other auxiliary electrical and mechanical equipment.
HIGH PRESSURE GAS SYSTEM.
The high pressure gas storage facility at the pad provides the launchvehicle with high pressure helium and nitrogen. This facility is
an il,:*o-ral part of the east portion of the launch support structure.It is entered from ground elevation on the east side of the pad.
The high pressure (o._,_'_3 psig) facilities at the pad are provided
for high pressure storage of 3,000 cubic feet of gaseous nitrogen
and 9,000 cubic feet of gaseous helium.
LAUNCH PAD INTERFACE STRUCTURE.
The launch pad interface structure (figure 7-10) provides mounting
support pedestals for the ML and MSS, an engine service platform
transporter, and support structures for fueling, pneumatic, electric
power and environment control interfaces.
The ML at the launch pad (as well as the VAB and refurbish
area) is supported by six mount mechanisms which are designed
to carry vertical and horizontal loading. Four extensible columns,located near each corner of the launcher base exhaust chamber,
LH 2 AND
GH 2 VENT
AND I
ENGINE
SATURN V
RP-1 MAST
SIDE 4
MECHANISM
(6 pLACES)
SIDE 1
SIDE 3
POWER
:ACILITIES
STAIRWAY
LOX
C.H 20047
Figure 7-10
7-1 1
Section VII Ground Support Interface
also support the ML at the launch site. These columns are designedto prevent excessive deflections of the launcher base when thevehicle is fueled and from load reversal in case of an abort between
engine ignition and vehicle liftoff.
The MSS is supported on the launch pad by four mounting mecha-
nisms similar to those used to support the ML.
The engine servicing platform provides access to the pedestal deck
area for servicing of the S-IB engines.
Interface structures are provided on the east and west portions
of the pad structure (figure 7-10) for propellant, pneumatic, power,facilities, environmental control, communications, control and in-
strumentation systems.
Modification for Skylab Saturn IB consists of re-routing the RP-1
supply line interface to the LH 2 mast and deactivating the RP-Imast used for Saturn V.
EMERGENCY INGRESS/EGRESS
AND ESCAPE SYSTEM.
The emergency ingress/egress and escape system (figure 7-11)
provides access to and from the Command Module (CM) plus an
escape route and safe quarters for the astronauts and service per-
sonnel in the event of a serious malfunction prior to launch. De-
pending upon the time available, the system provides escape byeither slide wire or elevator. Both means utilize the CM access
arm as a component.
The slide wire egress system (figure 7-11) provides the primary
means of escape. A 1 l/8-inch diameter steel cable extends from
the 341.72-foot level of the ML to a tail tower approximately 2,200feet west of the ML. Astronauts and technicians evacuate the white
room, cross the access arm, and follow a catwalk along the east
and north sides of the ML to the egress platform at the 320-foot
level. Here, they board the 9-man cab transporter suspended from
the cable and snubbed against the egress platform. The cab isreleased by levers in the cab. It rides the slide wire down to the
landing area where it is decelerated and stopped by an arresting
gear assembly.
The secondary escape and normal egress means are the tower high
speed elevators. These move between the 340 foot level of the
tower and level A at 600 feet per minute. At level A, egressing
personnel move through a vestibule to elevator No. 2 which takes
them down to the bottom of the pad. Armored personnel carriers
are available at this point to remove them from the pad area.
When the state of the emergency allows no time for retreat by
motor vehicle, egressing personnel upon reaching level A of the
ML slide down the escape tube into the rubber-lined blast room
vestibule (figure 7-11). The escape tube consists of a short sectionwhich extends from the elevator vestibule at ML level A to side
3 of the ML base where it interfaces with a fixed portion that
penetrates the pad at an elevation of 48 feet. At the lower extremity
of the illuminated escape tube, a deceleration ramp is provided
to reduce exit velocity, permitting safe exit for the user.
Entrance to the blast room is gained through blast-proof doorscontrollable from either side. The blast room floor is mounted on
coil springs to reduce outside acceleration forces to 3 to 5 G's. Twen-
ty people may be accommodated for 24 hours. Communication facil-
ities are provided in the room including an emergency RF link
in which the receiving antenna is built into the ceiling. In the event
that escape via the blast-proof doors is not possible, a hatch in
the top of the blast room is accessible to rescue crews.
An underground air duct from the vicinity of the blast room tothe remote air intake facility permits egress from the pad structure
to the pad perimeter. Provision is made to decrease air velocity
in the duct to allow personnel movement through the duct.
7-12
Emergency ingress to the CM utilizes the tower high speed elevatorsand the CM access arm.
ELECTRICAL POWER.
The electrical power for the launch pad is fed from the 69 kv
main substation to switching station No. 1, where it is stepped
down to 13.8 kv. The 13.8 kv power is fed to switching stationNo. 2 from where it is distributed to the various substations in
the pad area. The output of each of the substations is 480 volts
with the exception of the 4160-volt substations supplying power
to the fire protection water booster pump motors and the lox pumpmotors.
FUEL SYSTEM FACILITIES.
The fuel facilities, located in the northeast quadrant of the pad
approximately 1,450 feet from pad center, store RP-I and liquid
hydrogen.
The RP-I facility consists of three 86,000-gallon (577,000-pound)
steel storage tanks, a pump house, a circtilating pump, a transfer
pump, two filter-separators, an 8-inch stainless steel transfer line,
RP-I foam generating building and necessary valves, piping, and
controls. Two concrete RP-I holding ponds, 150 feet by 250 feet
with a water depth of two feet, are located north of the launch
pad, one on each side of the north-south axis. The ponds collect
spilled RP-1 and water drainage from the pad area. Traps permit
outflow of water from the holding ponds, but retain the fuel for
skimming and disposal.
The LH 2 facility consists of one 850,000-gallon spherical storage
tank, a vaporizer/heat exchanger which is used to pressurize the
storage tank to 65 psig, a vacuum-jacketed, 10-inch. Invar transfer
line and a burn pond venting system. The internal tank pressure,
maintained by circulating LH 2 from the tank through the vaporizer
and back into the tank, is sufficient to provide the proper flow
of LH 2 from the storage tank to the vehicle without using a transfer
pump. Liquid hydrogen boiloff from the storage and ML areasis directed through vent piping to bubblecapped headers submerged
in the burn pond. The hydrogen is bubbled to the surface of the
100-foot square, water filled, concrete pond where a hot wire igni-
tion system maintains the burning process.
LOX SYSTEM FACILITY.
The lox facility is located in the northwest quadrant of the pad
area, approximately 1,450 feet from the center of the pad. The
facility consists of one 900,000-gallon spherical storage tank, a lox
vaporizer to pressurize the storage tank, 1000 gpm fill pumps, a
vacuum jacketed transfer line, and a 10 in. uninsulated vent linefrom lut to drain basin.
Pad B lox system modifications for Skylab Saturn IB consist of
using the 1000 gpm pumps and vacuum-jacketed line for fill and
replenish; and disabling the un-insulated 10,000 gpm main lox
fill pumps and line used for Saturn V.
GASEOUS HYDROGEN FACILITY.
This facility is located on the pad perimeter road northwest of
the liquid hydrogen facility. The facility provides GH 2 at 6,000pslg to the launch vehicle. The facility consists of four storage tanks
having a total capacity of 800 cubic feet, a flatbed trailer on which
are mounted liquid hydrogen tanks and a liquid-to-gas converter,
a transfer line and necessary valves and piping.
AZIMUTH ALIGNMENT BUILDING.
The azimuth ahgnment building is located in the approach ramp
to the launch structure in the median of the crawlerway about
700 feet from the ML positioning pedestals. The building houses
Section VII Ground Support Interface
RUBBER ROOM
EGRESS rUNNEL
(AIR INTAKE DUC;)_-_ ...... I
J'--" ECS
;'RI_'_EBLAST ROOM EGRESS SYSTEM
ML
II s
EGRESS
STATION _
__ _/--___2_'_± _
E WIRE EGRESS SYSTEM
C-H 20048
Figure 7-11 7-13
Section VII Ground Support Interface
the auto-collimator theodolite which senses, by a light source, the
rotational output of the stable platform. A short pedestal, with
a spread footing isolated from the building, provides the mountingsurface for the theodolite.
PHOTOGRAPHIC FACILITIES.
Modifications have been made for Skylab Saturn IB to adjust thewater flow to the ML, to reduce the flow to the flame deflector,
and to delete excess storage tanks.
AIR INTAKE BUILDING.
This building houses fans and filters for the air supply to the PTCR,pad cellular structure and the ML base. The building is located
west of the pad, adjacent to the perimeter road.
FLAME DEFLECTOR.
The flame deflector is positioned under the vehicle in the flametrench and routes the booster exhaust away from the pad complex.
These facilities support photographic camera and closed circuit
television equipment to provide real-time viewing and photographic
documentation coverage. There are six camera sites in the launch
pad area, each site containing an access road, five concrete camera
pads, a target pole, communication boxes and a power transformer
with a distribution panel and power boxes. These sites cover pre-
launch activities and launch operations from six different angles
at a radial distance of approximately 1,300 feet from the launch
vehicle. Each site has four engineering sequential cameras and one
fixed, high speed, metric camera (CZR). A target pole for optical
alignment of the CZR camera is located approximately 225 feet
from the CZR pad and is approximately 86 feet high.
PAD WATER SYSTEM FACILITIES.
The pad water system facilities supply water to the launch pad
area for fire protection, cooling, and quenching. Specifically, the
system furnishes water for the industrial water system, flame deflec-
tor cooling and quench, ML and pedestal deck cooling and quench,
ML tower fogging and service arm quench, sewage treatment plant,Firex water system, lox and fuel facilities, ML and MSS fire protec-
tion and all fire hydrants in the pad area. The water is suppliedfrom three 6-inch wells, each 275 feet deep. The water is pumped
from the wells through a desanding filter and into a 1,000,000-gallonreservoir.
MOBILE SERVICE STRUCTURE (MSS).
The mobile service structure (figure 7-12) provides access to those
portions of the space vehicle which cannot be serviced from theML while at the launch pad. During nonlaunch periods, the MSS
is located in a parked position along side of the crawlerway, 7,000feet from the nearest launch pad. The MSS is transported to the
launch site by the C-T. It is removed from the pad a few hours
prior to launch and returned to its parking area.
The MSS is approximately 402 feet high, measured from ground
level, and weighs 12 million pounds. The tower structure rests ona base 135 feet by 135 feet. The top of the MSS base is 47 feet
above grade. At the top, the tower is 87 feet by 113 feet.
The MSS is equipped with systems for air conditioning, electrical
power, communications networks, fire protection, nitrogen pres-surization, hydraulic pressure, potable water and spacecraft fueling.
The structure contains five work platforms which provide access
to the space vehicle. The outboard sections of the platforms are
actuated by hydraulic cylinders to open and accept the vehicleand to close around it to provide access to the launch vehicle and
7-14
spacecraft. The three upper platforms are fixed but can be relocated
as a unit to meet changing vehicle configurations. The uppermost
platform is open, with a chain-link fence for safety. The two plat-
forms immediately below are enclosed to provide environmental
control to the spacecraft. The two lowest platforms can be adjusted
vertically to =serve different parts of the vehicle. Like the uppermost
platform, they are open with a chain-link fence for safety.
Platform No. 2 is modified for the exterior configuration of the
Saturn IB vehicle, and Platform No. I is modified for the Skylab
Saturn V vehicle. After Skylab 1 and 2 launches, Platform No.
1 will be modified to be compatible with the Saturn IB vehicle.
CRAWLER-TRANSPORTER (C-T).
The crawler-transporter (figure 7-13) is used to transport the mobile
launcher and the mobile service structure. The ML, with the space
vehicle, is" transported from the vehicle assembly building to the
launch pad. The MSS is transported from its parking area to and
from the launch pad. After launch, the ML is taken to the refurbish-
ment area and subsequently back to the VAB. The C-T is capable
of lifting, transporting, and lowering the ML or the MSS withoutthe aid of auxiliary equipment. The C-T supplies limited electric
power to the ML and the MSS during transit.
The C-T consists of a rectangular chassis which is supported through
a suspension system by four dual-tread crawler-trucks. The overall
length is 131 feet and the overall width is 114 feet. The unit weighs
approximately 6 million pounds. The C-T is powered by self-con-
tained, diesel-electric generator units. Electric motors in thecrawler-trucks propel the vehicle. Electric motor-driven pumps
provide hydraulic power for steering and suspension control. Air
conditioning and ventilation are provided where required.
The C-T can be operated with equal facility in either direction.Control cabs are located at each end and their control function
depends on the direction of travel. The leading cab, in the direction
of travel, will have complete control of the vehicle. The rear cab
will, however, have override controls for the rear trucks only.
Maximum C-T unloaded speed is 2 mph, 1 mph with full load
on level grade, and 0.5 mph with full load on a five percent grade.
It has a 500-foot minimum turning radius and can position the
ML or the MSS on the facility support pedestals within ± two
inches.
CONVERTER/COMPRESSORFACILITY.
The primary gaseous nitrogen source for LC-39 is via a cross country
line from an industrial supplier; however, the on-site Converter/
Compressor Facility (CCF) operates as a back-up source. For
gaseous heliiam the CCF compresses helium brought in by trucksand railroad cars. The CCF is located in tlae northside of the
crawlerway between the VAB and the MSS park area. The CCF
converts liquid nitrogen to low pressure and high pressure gaseous
nitrogen and compresses gaseous helium. It thegn supplies the gases
to the storage facilities at the launch pad and at the VAB.
The facility includes a 500,000-gallon storage tank for liquid ni-
trogen, tank vaporizers, high pressure liquid nitrogen pump and
vaporizer units, high pressure helium compressor units, helium and
nitrogen gas driver/purifiers, rail and truck transfer facilities anda data link transmission cable tunnel.
After processing, the gaseous nitrogen is piped to the distribution
lines supplying the VAB area (6,000 psig) and the pad (150 psig
and 6,000 psig). The gaseous helium is stored in tube-bank rail
cars.TheheliumpassesthroughtheCCFheliumcompressorswhichboostitspressurefromthetube-bankstoragepressureto6,000
Section VII Ground Support Interface
psig after which it is piped to the VAB and pad high pressure
storage batteries. Mass flow rates of high pressure helium, high
MOBILE SERVICE STRUCTURE
NOTES:
PLATRDRM NO. 1 WILL NOT BE USED FOR C/O OF
/_k SL-2 BUT AFTER MODIFICATION WILL BE USED FOR
C/O OF SL-3 & 4.
PLATFORM NO. 2, BY USE OF (4) NEW PORTABLE
/_ ACCESS PLATFORMS, WILL BE USED FOR C/O OFSL-2.
THE (4) NEW PORTABLE ACCESS PLATFORMS DESIGNED
/_ FOR pLATFORM NO. 2 WILL BE MOVED TO PLAT FORM
NO. 1 FOR CHECKOUT OF SL-3 & 4.
(2) ACCESS PLATFORMS WILL BE ADDED TO PLATFORM
/_ NO. 3A FOR ACCESS TO THE IU COMMAND ANTENNAS.
355'-0" LEVEL-
313'-½"LVL-H-
221'-10½"
177'-6"
88'-9" LEVEL
91'-2"LVL-C -
44'-4½" LEVEL
ORERATIONS
SUPPORT
ACE
MOBILE SERVICE STRUCTURE
SKYLAB-2, 3&4
Figure 7-12
-PLATFORM NO. 5
II "
rPLATFOpdv_ NO. 3B
• PLATFORM NO. 3A
/xPLATFORM NO.
S/A-6
MOBILE LAUNCHER NO. I
-+340' LEVEL
,320' _yEL
• 300' LEVEL
-. 280' LE_VEL
260' LEVEL
•240' LEVEL
- 220' LEVEL
- 200' LEVEL
- 180' LEVEL
-160' LEVEL
- 140' LEVEL
- 127' DECK LEVEL"
- 120' LEVEL
100' LEVEL
80' LEVEL
- 60' LEVEL
30' LEVEL
- 0 LEVEL
--A LEVEL
LEVEL
C.H 20049
7-15
Section VII Ground Support Interface
pressure nitrogen, and low pressure nitrogen gases leaving the CCFare monitored on panels located in the CCF.
ORDNANCE STORAGE AREA.
The ordnance storage area serves LC-39 in the capacity of labora-
tory test area and storage area for ordnance items. This facility
is located on the north side of the crawlerway and approximately2,500 feet northeast of the VAB. This remote site was selected
for maximum safety.
The ordnance storage installation, enclosed by a perimeter fence,
is comprised of three archtype magazines, two storage buildings,
one ready-storage building, an ordnance test building and a guardservice building. These buildings, constructed of reinforced concrete
and concrete blocks, are over-burdened where required. The facility
contains approximately 10,000 square feet of environmentally con-
trolled space. It provides for storage and maintenance of
retrorockets, ullage rockets, explosive separation devices, escape
rockets and destruct packages. It also includes an area to test the
electro-explosive devices that are used to initiate or detonate ord-
nance items. A service road from this facility connects to SaturnCauseway.
VEHICLE ASSEMBLY ANDCHECKOUT.
The vehicle stages and the instrument unit (IU) are, upon arrivalat KSC, transported to the VAB by special carriers.
The S-IB stage is erected on hold-down arms installed in the transfer
aisle for installation of fins then is lifted onto the pedestal of ML-I
in High Bay 1. The S-IVB stage and the IU are delivered to
preparation and checkout cells in the low bay area for inspection,
checkout, and pre-erection preparations.
The S-IVB stage and IU are then moved to the transfer isle and
over the diaphram into High Bay 1 and assembled vertically on
the ML with the S-IB stage. The spacecraft and launch escapesystem are then assembled.
Following assembly, the space Vehicle is connected to the LCC
via a high speed data link for integrated checkout and a simulated
flight test. When checkout is completed, the crawler-transporter
(C-T) picks up the ML, with the assembled space vehicle, and
moves it to the launch site over the crawlerway.
7-16
Figure 7-13
For the initial Skylab Saturn IB mission, SL-2, a boilerplate space-
craft (BP-30) is installed and the vehicle is moved to the Pad forfacilities checkout. The vehicle is then returned to the VAB for
installation of the flight spacecraft.
At the launch site, the ML is emplaced and connected to system
interfaces for final vehicle checkout and launch monitoring. The
mobile service structure (MSS) is transported from its parking area
by the C-T and positioned on the side of the vehicle opposite the
ML. A flame deflector is moved on its track to its position beneath
the blast opening of the ML to deflect the blast from the S-IB
stage engines.
For launch of the SL-I, the MSS will be moved briefly to Pad
A for inspection of the vehicle, then returned to Pad B.
During the prelaunch checkout, the final system checks are com-
pleted, the MSS is removed to the parking area, propellants are
loaded, various items of support equipment are removed from the
ML and the vehicle is readied for launch. After vehicle launch,
the C-T transports the ML to the VAB for refurbishment.
TEST SYSTEM.
The requirement to launch Saturn IB vehicles from the Saturn
V launch site mandated a reconfiguration of LC-39-3 (FR-3, ML-I,
Pad B, and High Bay 1). The launch vehicle ground support equip-
ment (LVGSE) being used to check out and launch the Skylab
Saturn IB vehicles is basically Saturn V equipment located atLC-39-3 supplemented by hardware relocated from LC-34/37B and
by some new equipment. The S-IB ESE, for example, utilizes
converted S-II ESE, augmented by S-IC ESE and some LC-34/37B
hardware. The Saturn V ML-3 Terminal Countdown Sequencer
(TCS) will be used for the Saturn IB automatic sequence. MGSE
systems will utilize Saturn V MGSE, as requlred and the S-IBstage calips console relocated from LC-37B. The LC-39-110A
ground computer and display system will be utilized for Saturn
IB, but new software programs will be prepared.
A computer controlled automatic checkout system is used to ac-
complish the VAB (high bay) and pad testing. A I10A ground
computer and the equipment necessary to service and check out
the launch vehicle are installed on the ML. Also a I IOA ground
computer and the display and control equipment necessary to
monitor and control the service and checkout operations are in-
stalled in the LCC. The computers operate in tandem through adata link. The computer in the ML receives commands from and
transmits data to the computer in the LCC. The physical arrange-
ment of the LCC and the ML are illustrated in figures 7-14 and7-5 respectively.
Test System Operation.
Test system operation for Saturn IB launch vehicle checkout is
conducted from the firing room (see figure 7-15). During prelaunch
operations, each stage is checked out utilizing the stage control
and display console. Each test signal is processed through the
computer complex, and is sent to the vehicle. The response signal
is sent from the vehicle, through the computer complex, and the
result is monitored on the display console. The basic elements of
the test system and their functional relationship are shown in figure7-16.
A switch on the control console can initiate individual operationof a system component or call up a complete test routine from
the computer. The insertion of a plastic coded card key, prior to
console operation, is a required precaution against improper pro-
gram callup. Instructions, interruptions and requests for displays
are entered into the system by keying in proper commands at theconsole keyboards.
A complete test routine is called up by initiating a signal at the
control panel. The signal is sent to the patch distributor located
Section VII Ground Support Interface
COMPUTER
DISPLAY ROOM
3RD FLOOR
FIRING ROOM
DISPLAY SCREENS
LAUNCH MANAGEMENT TEAM
2ND FLOOR
TELEMETRY AND RF
1ST FLOOR
OFFICES, CAFETERIA AND DISPENSARY
Figure 7-14
C-H 20051
7-17
Section VII Ground Support Interface
FIRING RO( M TYPICAL)
_W
m
1 -
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LVD
DOCUMENTATION
CENTER
SHAFT
rl [I-'_L__ SIG COND
INTEGRATED/
Ill l II III III
1.4._ PATCHOts'r.------_
II Illlll II
•=_ MEAS REC---,_J
illlllililll II III11 lii,e-------- MEAS REC-_---_J
I I III I I II II I III I II II I II I I--'FATCH_CKS/
_AS REC------*I r"_NETWORKS_*I i_ -- q= Drs--- 1IIIII IIIIIII I I I I II II I I IIIII IIII,_-----MEAS_C_ J*--AIRCON°EQUIF--_I p---- FATCHRACKS_=Illlllillllll i I I I r-l-l-1 []
Jl MEAS & RF I=J
III I111111 IIIIIII i I I I I I I I I
FLIGHT CONTROL I,J
IIIIIIIIIIIIIIII! I l i I I I I I
J4_ STABILIZATION _ GUI DANCE-'_J
Ililllll li IlllllI I I I I I I I I
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NOTES:
I. LVO DEP. DIR.
FOR ENGIN.
2. MSFC PROGRAM
MANAGER
3. LVD DIR.
4. LVO
5. LAUNCH DIR.
6. KSC DIR,
7. DEP. DIR.
LAUNCH OPNS.
8. DIR. SCO
9. MSC SKYLAB
FROG. MGR.
I0. PUBLIC
AFFAIRS
OFFICER
MEN
ROOM
1C-H 200S2
7-18 Figure 7-15
"1"1
(D
(3)
CO
PAD AND TOWER
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Section VII Ground Support Interface
in the LCC and is routed to the appropriate signal conditioningequipment where the signal is prepared for acceptance by the LCC
computer complex. The LCC computer communicates with the ML
computer to call up the test routine. The ML computer complex
sends the signal to the ML signal conditioning equipment and then
to the stage relay rack equipment. The signal is then routed to
the terminal distribution equipment and through the crossover
distributor to interrogate the vehicle sensors. The sensor outputs
are sent back to the ML computer complex for evaluation. The
result is then sent to the LCC computer complex which routes
the result to the stage console for display. Manual control of vehicle
functions is provided at the control consoles. This control bypasses
the computers and is sent to the vehicle by means of hardwire.
The result is also sent back to the display console by hardwire.
The digital data acquisition system (DDAS) collects the vehicle
and support equipment responses to test commands, formats thetest data for transmission to the ML and LCC, and decommutates
the data for display in the LCC. Decommutated test data is fed
to the LCC computers for processing and display, and for computer
control of vehicle checkout. The DDAS consists of telemetry equip-
ment, data transmission equipment and ground receiving stations
to perform data commutation, data transmission and data decom-mutation.
The digital event evaluators (DEE) are used to monitor the status
of input lines and generate a time tagged printout for each detected
change in input status. High speed printers in the LCC are con-
nected to each DEE to provide a means for real-time or post-test
evaluation of discrete data. Two systems (DEE-3 and DEE-6) areused to monitor discrete events.
Two DEE-3's are located in the PTCR with a printer located in
the LCC. One unit monitors 768 inputs and the other monitors
1032 inputs, all associated with propellant loading, environmentalcontrol, and water control.
The DEE-6 processor, offtine printer and remote control panel are
located in the LCC, while interface equipment is located in the
ML base. Several remote printers are located strategically within
the firing room. The DEE-6 monitors up to 4320 discrete signals
from the vehicle stage umbilicals, pad and tower ground supportequipment.
The computer complex consists of two 110A general purpose com-
puters and peripheral equipment. This equipment includes a line
printer, card reader, card punch, paper tape reader and magnetic
tape transports. The peripheral equipment provides additional bulk
storage for the computer, acts as an input device for loading test
routines into the computer memory and as an output device to
record processed data. One computer is located in the ML base
and the other in the LCC. The computers are connected by un-
derground hardwire. The LCC computer, the main control for the
system accepts control inputs from test personnel at the consoles
in the firing room as well as inputs from tape storage and transmits
them as test commands to the ML computer. The ML computer
has the test routines stored in its memory banks. These routines
are called into working memory and sent as discrete signals to
the launch vehicle in response to the commands received from
the LCC computer. The ML computer reports test routine status,
data responses and results of test to the LCC computer. It is through
this link that the control equipment and personnel in the firing
room are informed of the test progress.
The propellant tanking computer system (PTCS) determines and
controls the quantities of fuel and oxidizer on board each stage.
Optimum propellant levels are maintained and lox and LH 2 are
replenished as boiloffoccurs during the countdown. The propellant
tanking operation is monitored on the PTCS control panel.
The propellant Data Transmission System (DST) is the digital
communication link between the LCC and the PTCR for propellant
MAJON fVENT:
7-20
loading and environmental control command and monitor data.
The system operates on a time-shared mode with data transmitted
over video pairs. Each input signal is sampled about once a second.
Visual surveillance of launch vehicle checkout is provided to the
launch management team and for distribution to MSC and MSFC
through the operational television system (OTV). Fifty two cameras
provide this capability, 18 of which are located on the ML, in
the pad area, 12 on the MSS and 6 in the LCC. Any camera
may be requested for viewing on the 10 x 10 foot screens in the
firing room.
Vehicle and ground instrumentation measurements may be mom-
tored by firing room personnel using the television data display
system (TDDS). Stage DDAS, ground DDAS, and facility and
environmental measurements are processed in the CIF telemetry
ground station, and prepared for computer entry. The CIF GE-635
scientific computer is used to process these measurements to provide
engineering units to be displayed on I0 dual eight-inch Operational
Television (OTV) monitors in the firing room. The computer canalso check each measurement for out-of-limit conditions, and alert
the operator either visually or audibly if pre-selected limits are
exceeded. The operator has the option of selecting pre-defined pages
or operator defined pages which contain up to sixteen measure-
ments, pre-defined sets of measurements where only those that
have out-of-limit conditions are displayed, or a graphical presenta-
tion of an individual measurement. The graph can be generated
in real time, or from the last 51 minutes of history, or a combinationof both.
Certain major events may be observed by members of the launch
management team who occupy the first four rows in the firing
room. The significant launch vehicle events which are displayed
on the 10 x 10 foot screen are shown in figure 7-17.
PRELAUNCH OPERATIONS.
The prelaunch operations (figure 7-18) take place in the Manned
Spacecraft Operations (MSO) Building, the VAB and the launch
pad.
MSO Building Activities.
After receipt of the spacecraft stages at KSC, inspection, testing,
assembly and integrated checkout of the spacecraft take place in
the MSO building. The assembled spacecraft is transported to theVAB and mated to the launch vehicle.
LAUNCH S-IB ONA5-206 SEQUENCE START INTERNAL POWER
RANGE S-IVB LOX TANK
SAFE PRESSURIZED
S-IVB LH 2 TANKPRESSURIZED
S-IVB PROPELLANTS
PRESSURIZED
_UNCH SUPPORT S-IB FUEL TANKS LSE READY
PREPS COMPLETE. PRESSURIZED FOR IGNITION
S-IVB PREP S-IB LOX TANKS READY FORCOMPLETE PRESSURIZED S-18 IGNITION
IU READY S-IB PROPELLANTS START S-18PRESSURIZED IGNITION
ED5 READY AUDIO SE_. POSIT COMMITNO. 3 INDICATION
5-1B PREP AUDIO SEL POSITCOMPLETE NO. 4 INCICATIO N LIFTOFF
Figure 7-17
S-)B STAGE
CUTOFF
51B/S-IVB SEP
LOGIC ZERO
S-IVB ENGINE
START
S-IV8 CUTOFF
TEST
HOLDING
TEST
COUNT(NG
EVE i'_q"SYSIE_
CALf GRATING
CH._OS4
VAB Activities.
The VAB activities are the assembly and checkout activities which
are completed in two major areas of the VAB; the high bay and
the low bay.
Low Bay Activities. The low bay activities include receipt and
inspection of the S-IVB stage and IU and the assembly and checkout
of the S-IVB stage.
The S-IVB stage is brought into the low bay area and positioned
on the checkout dolly. A fuel tank inspection, J-2 engine leak test,
hydraulic system check and propellant level sensor electrical checksare made.
High Bay Activities. High bay activities include S-IB stage checkout,stage mating, stage systems tests, launch vehicle integrated tests,
space vehicle overall tests, and a simulated flight test. High bay
checkout activities are accomplished using the consoles in the firing
room, the computer complex, and display equipment.
The S-IB stage is positioned and secured to the ML and access
platforms are installed: The umbilicals are secured to the vehicle
plates. Prepower and power-on checks are made to ensure electrical
continuity, Pneumatic, fuel, lox and H-I engine leak checks aremade. Instrumentation, and range safety system checks are made.
The S-IVB stage is mated to the S-IB stage and the IU is matedto the S-IVB. The S-IVB and IU umbilicals are secured to the
vehicle plates. Pre-power and power-on checks are made to ensure
electrical continuity. S-IVB engine hydraulic, pressurization and
auxiliary propulsion system leak checks are made. S-IVB propellant,
propulsion, pressurization and range safety system checks are made.
1U S-band, and C-band and guidance and navigation system checksare made.
Following completion of the stage system tests, launch vehicle
integrated checks are accomplished. Vehicle separation, flight con-
trol, sequence malfunction and emergency detection system checks
are made. The spacecraft is then mated to the launch vehicle.
OPERATIONS COMPLETE(DAYS BEFORE LAUNCH)
CSM OPERATIONS
COMBINED SYSTEM TEST
UNMANNED ALTITUDE RUN
MANNED ALTITUDE RUN
CSM/SLA MATEORDNANCE INSTALLATION
LV VAB HIGH BAy OPERATIONS
S-IB ERECTION
LV ERECTION
LV ELECTRICAL SYSTEM TEST
LV MALFUNCTION OVERALL TEST
LV SERVlCE ARM OVERALL TEST
SPACECRAFT ERECTION
SPACE VEHICLE VAB OPERATIONS
TRANSFER TO PAD
PAD OPERATIONS
LV POWE R ON
SV PLUGS iN TEST
SV FLIGHT ELECTRICAL MATING
SV BACKUP GUIDANCE TEST
SV FLIGHT READINESS TEST
SV HYPERGOLtC LOADING
S-IB RP-I LOADING
CDDT-WET/DRY
COUNTDOWN
L-228
L-130
L-110
L-100
L-98
L-242
L-235
L-213
L- 173
L-131
L-81
L-71
L-69
L-67
L-41
L-40
L-38
L-34
L-21
L-19
L-10
L-0
Figure 7-18
Section Vll Ground Support Interface
For the first checkout flow with SA-206, a boilerplate spacecraft
(BP-30) is installed and the vehicle is moved to the pad for facilitiescheckout. The vehicle is then returned to the VAB for installation
of the flight spacecraft.
After the spacecraft is mated, two space vehicle tests are made.
Test number 1 is performed to verify RF, ordnance, pressurization,
propulsion, guidance and control, propellant and emergency detec-
tion system operation. Test number 2 is performed to verify proper
operation of all systems during an automatic firing sequence and
flight sequence. This includes a simulated holddown arm release,
electrical umbilical ejection, swing arm retraction and firing of live
ordnance in test chambers; and a simulated flight sequence that
verifies proper operation of the space vehicle during a normal minus
count and an accelerated plus count. A normal mission profile
is followed during this time. The simulated flight test ensures that
the space vehicle is ready for transfer to the pad. The launch escape
system is installed on the command module of the spacecraft. TheML and space vehicle are now ready for transport to the pad.
Vehicle Transfer and Pad Mating Activities.
After completion of the VAB activities, the ML transports the
assembled space vehicle to the launch pad. Approximately eight
hours are required for this operation. The space vehicle/ML are
then interfaced with the launch pad.
Pad Activities.
For the facility checkout flow of the pad, SA-206 and boilerplatespacecraft BP-30 will be moved to the pad. Facility checkout activi-
ties will include an LV cryogenic loading sequence; a LUT/PAD
water system test; and an RP-I loading sequence. After facility
checkout is completed, the vehicle is returned to the VAB for
installation of the flight spacecraft. The space vehicle will then
be moved to the pad for prelaunch checkout and servicing.
In general, once the vehicle and ML have been mated to the pad
facility, two major operations must be performed. The first is to
verify the readiness of the launch vehicle, spacecraft and launch
facility to perform the launch sequence and the second is to com-
plete the launch operation.
The Countdown Demonstration Test (CDDT) verifies that the
launch vehicle and the ground support equipment are in launch
status. The CDDT is performed in two phases, the wet CDDT
and the dry CDDT.
The wet CDDT is performed the same as the launch countdown
with the following major exceptions:
1. Service arms are pressure tested.
2. Digital range safety command system test code plugs are used
instead of flight code plugs.
3. Hypergol cartridges, igniters, initiators, safe and arm devices,
and exploding bridgewire detonators are inert.
4. Astronauts do not board the spacecraft.
5. Terminal count sequence is interrupted at time for ignition.
The dry CDDT is performed the same as the last 3 1/2 hours of
the launch countdown with the following major exceptions:
1. Launch vehicle cryogenic propellants are not on board.
2. The primary damper is not disconnected.
3. Service Arm No. 9 is reconnected as soon as the system has
stabilized in the park position.
4. Hypergol cartridges, igniters, initiators, safe and arm devices.
and exploding bridgewire detonators are inert.
5. Terminal count sequence is not initiated.
7-21
Section VII Ground Support Interface
Following the CDDT, preparations for the actual countdown are
started. The preparations include items which would either compro-mise the safety of the vehicle if done later in the countdown or
impose additional constraints on pad access during the final phasesof '.he countdown.
Approximately six days before the launch readiness day, the count-
down begins and the space vehicle is subjected to the final checkoutand servicing operations required for launch.
Concurrently, the Skylab 1 space vehicle is undergoing prelaunchcheckout and launch operations on Pad A. The countdown for
Skylab 2 is timed such that the countdown clock will be started
(at T-22:30 hours) upon lift-off of SL-I.
The final phase of the countdown starts approximately nine hours
prior to liftoff. During the final phase, the cryogenics are loaded,
conditioned and pressurized. Final checks are made on all subsys-
tems. The propulsion systems are serviced and prepared for launch.
All onboard spheres are brought up to flight pressure and the crewmans the Command Module.
By the time spacecraft closeout is complete, most major operations
have been completed. Propellants are being replenished as requiredto supplement cryogens lost due to boiloff. Boiloff will continue
until the various stage vent valves are closed for tank prepressuriza-tion and some vapor may be noticeable.
With the start of the automatic sequence at T-187 seconds, the
final operations required for launch begin. All pneumatic and
propellant supply lines are vented and purged to prevent damageto the vehicle at umbilical release. The vehicle is switched to internal
power, necessary purges are put in launch mode and some servicesare retracted.
At T-3 seconds, the S-IB ignition command is given. At T-0 seconds,
the launch commit signal is given, causing the eight holddownarms to retract. These arms restrain the launch vehicle until a
satisfactory thrust level is achieved after which the controlled release
assemblies provide for gradual release of the vehicle during liftoff.
HOLD AND RECYCLE CRITERIA.
Interruption of the countdown due to equipment failure, weatheror other causes may occur at any time. When the countdown is
interrupted, subsequent feasible actions depend on the function
taking place at the time. These actions include holding and/or
recycling. Feasible actions also, in some cases, are affected byprevious operations conducted on the vehicle, such as the number
of pressure cycles the propellant tanks have undergone.
A hold is defined as an interruption of the countdown for unfavor-
able weather, repair of hardware, or correction of conditions un-
satisfactory for launch or flight.
In a recycle, the countdown is stopped and returned to a designated
point as specified in the launch mission rules.
For a scrub, the launch attempt must be rescheduled for a laterwindow.
A turnaround comprises the actions required to recycle, hold until
the countdown can be resumed for a specific window, and complete
the countdown from the re-entry point to T-0.
Decision/repair time is the time available to make decisions and/or
conduct repair operations before initiating the count.
LAUNCH CONSTRAINTS.
Various operational, launch vehicle, spacecraft and support equip-
ment factors affect hold/recycle processing actions. Several of these
factors are briefly discussed in the following paragraphs.
7-22
The length of the launch window on any launch date is a mission
peculiar constraint. This constraint determines the maximum hold
limit for the countdown period between start of cryogenic loading
and commence S-IVB stage start bottle chilldown.
Launch vehicle batteries have a life of 120 hours following activa-tion. The batteries are installed in the vehicle at T-53.5 hours and
are activated 33.0 hours prior to their installation. Assuming a
countdown programmed hold of 15 hours, the battery life expendedin a normal count is 88.0 hours, leaving an available battery lifeof 32.0 hours.
The Safe and Arm (S&A) devices are remotely controlled ordnance
items used to make safe and/or arm the launch vehicle propellant
dispersion systems. The devices are certified at T-108 hours. Recer-
tification is required in seven calendar days. Device removal from
the vehicle is required to perform recertification. The devices areinstalled at T-110 hours and connected at T-46.5 hours. In a normal
count, the allowable S&A device life is 59 hours.
The S-IVB Auxiliary Propulsion System (APS) modules are serviced
with hypergolic propellants. Pressurization of the system is done
at T-40 minutes. Depressurization and gas removal must be ac-
complished in the event of a scrub if access is required in the
S-IVB access control area. This task takes two hours and requiresuse of the Mobile Service Structure (MSS).
The Command and Service Module (CSM) fuel cell cryogenics
provide the electrical power for the spacecraft. The water resulting
from the reaction in the fuel cell is used for drinking purposes
during space missions. The cryogenic tanks are loaded to sufficient
capacity to tolerate a 56 hour delay for a normal count. Water
generated by fuel cell operation must be drained if a hold will
exceed 17 hours. Cryogenic replenishment is normally required
if turnaround exceeds 56 hours. The MSS is required for CSMcryogenic servicing.
Capacities of launch support facilities and equipment such as the
gaseous hydrogen and helium facilities, the cryogenic storage facili-
ties and the ground hydraulic supply unit affect hold and recycle
capability. For example, the gaseous hydrogen system can support
four complete cycles of start tank chilldown operations from T-22
minutes to T-4.5 minutes before recharging of the storage battery
is required. Recharging is accomplished by mobile units which
cannot be moved into position until launch vehicle cryogenics aredownloaded.
LAUNCH INTERLOCKS.
The terminal countdown sequencer (TCS) automatically controls
the space vehicle countdown beginning at T-187 seconds. Interlocks
are provided to prevent initiation of this sequence or to terminate
the sequence if conditions do not remain proper for launch. Once
the automatic sequence has been initiated, it can only be stopped
by a cutoff signal. There are no provisions for holding.
FIRING PREPARATION COMPLETE INTERLOCKS.
These interlocks insure that all vehicle and ground support equip-
ment systems are in proper condition for starting the automatic
sequence.
S-IB Stage ESE.
The following interlocks functions are provided by the S-IB stageESE.
a. Lox leak failure
b. No conax fired
c. Ignition armed
d. Safety switches armed
e. Flight sequence zero
f. Propellant dispersion system ready
g. Terminal countdown sequencer power supply OK
h. Spacecraft ready for launch
i. Instrument unit ready for launch
j. EDS ready
k. Range safety
I. Launch support equipment preparations complete
m. S-IVB systems ready
n. Prevalves open
o. Purges armed
p. Terminal countdown sequencer ready
q. + IDI61 (launch bus) supervision
S-IVB Stage ESE.
The following interlocks functions, which are summed as the "S-IVB
systems ready" interlock in the S-IB stage ESE, are provided by
the S-IVB stage ESE.
a. LH 2 tank wet
b. Non-programmed engine cutoff off
c. Passivation relays reset
d. S-1VB engine cutoff
e. Engine ready
f. Aux hydraulic pump flight mode relay not reset
g. Aux hydraulic pump power on
h. Ordnance OK
i. Ullage rocket pilot relays reset
j. APS No. l or No. 2 engine valve power on
k. Lox and LH 2 prevalves emergency close command off
1. Heat exchanger bypass valve enable relay reset
m. Mixture ratio control valve relays reset
n. Engine start relay reset
o. TM PI transmitter on
Instrument Unit ESE.
The following interlock functions, which are summed as "Instru-
ment Unit ready for launch" in the S-IB stage ESE, are provided
by the IU ESE.
a. LVDA firing commit inhibit A and B
b. IU guidance failure A and B
c. 56 volts OK
d. 400 Hz power A and B
e. Control voltage OK
f. Control attenuator timer at zero
g. Control accelerometer on
h. Gyro on
i. ST-124M system ready
j. Power transfer command and T-47
Section VII Ground Support Interface
k. IU power transfer complete
1. Switch selector internal power
m. Guidance alert and release
n. LVDA firin_t commit enable
EDS ESE.
The following interlock functions are provided by the EDS ESE
and are summed as the "EDS ready for launch" interlock in theS-IB ESE. i
a. EDS not ready
b. Ready circuits enable
READY FOR IGNITION INTERLOCKS.
These interlocks insure that all vehicle and ground support equip-
ment systems are in the proper condition for starting the S-IB engine
ignition s_quence. If the interlocks are not met at time for ignition,the countdown will terminate automatically. The count will then
have to be recycled to the start of automatic sequence. Once the
ignition sequence has started, a change of state of these functionswill not result in automatic termination of the countdown.
S-IB Stage ESE.
The following interlock functions are provided by the S-IB stageESE.
a. Lox leak failure
b. No conax fired
c. Ignition armed
d. Safety switches armed
e. Flight sequence zero
f. Propellant dispersion system ready
g. Terminal countdown sequencer power supply OK
h. Spacecraft ready for launch
i. Instrument unit ready for launch
j. EDS ready
k. Range safety
1. S-IVB stage ready for launch
m. Propellants pressurized
n. S-IB power transfer complete
o. Gas generator lox injector purge on
p. Thrust chamber fuel injector purge on
q. Launch bus (+ 1DI61) supervision
S-IVB Stage ESE.
The following interlock functions are provided by the S-1VB stageESE. These are summed as "S-IVB stage ready for launch" in
the S-IB stage ESE.
a. LH 2 tank wet
b. Non-programmed engine cutoff off
c. Passivation relays reset
d. S-IVB engine cutoff
e. Engine ready
f. Aux hydraulic pump flight mode relay not reset
7-23
SectionVIIGroundSupportInterfaceg. Aux hydraulic pump power on
h. Ordnance OK
i. Ullage rocket pilot relays reset
j. APS No. 1 or No. 2 engine valve power on
k. LH 2 directional vent in flight position
1. Power transfer complete
Instrument Unit ESE.
The following interlock functions are provided by the IU ESE.
These are summed as "IU ready for launch" in the S-IB stageESE.
a. LVDA firing commit inhibit A and B
b. IU guidance failure A and B
c. 56 volts OK
d. 400 Hz power A and B
e. Control voltage OK
f. Control attenuator timer at zero
g. Control accelerometer on
h. Gyro on
i. ST-124M systems ready
j. Power transfer command and T-47
k. IU power transfer complete
1. Switch selector internal power
m. Guidance alert and release
n. LVDA firing commit enable
EDS ESE.
The "EDS ready" interlock in the S-IB stage ESE is the "EDSnot ready" interlock in the EDS ESE. This interlock ensures that
the EDS control rate gyro rates are not excessive, enable logicis zero, cutoff enable I and 2 are not enabled, cutoff commands
1, 2, and 3 from the spacecraft have not been generated, and EDS
unsafe A and B have not been generated.
CUTOFF INTERLOCKS.
Automatic Sequence Start UnUl Tlme for IgniUon.
Cutoff interlocks for automatic sequence termination from the start
of automatic sequence until time for ignition are as follows:
a. Manual cutoff (switch function on S-IB networks panel)
b. Emergency manual cutoff(switch function on S-IB firing panel)
c. Premature ignition
d. Premature commit
e. Terminal countdown sequencer failure
f. Voltage failure (S-IB stage)
g. Sequence failure
Ignition Until Commit.
Fhe cutoff interlocks from ignition to commit are as follows:
a. Time for commit
b. All engines running
c. Absence of cutoff
Time for Ignition Until Commit.
The following functions are interlocked to provide cutoff betweentime for ignition and commit _f Conditions warrant.
a. Manual cutoff (switch function on S-IB networks panel)
b. Emergency manual cutoff (switch function on the S-IB firingpanel)
c. Terminal countdown sequencer failure
d. IU failure
e. EDS failure
f. Lox leak failure
g. Any engine cutoff
h. Thrust failure
Commit Until Plugs Separation.
The following functions are interlocked for cutoff between commit
and umbilical connector separation from the vehicle.
a. Manual cutoff (switch function on S-IB networks panel)
b. Emergency manual cutoff(switch function on S-IB firing panel)
c. Launch failure (liftoff has not occurred before launch failure
timers have expired)
7-24
SECTION VIII
I ..ss.o.co.. o, ITABLE OF CONTENTS
IntroducUon .................................................................... 8-1
Mission Control Center ............................................... ._ 8-1
Launch Control Center .................................................. 8-3
Abort Ground Rules ........................................................ 8-7
Huntsville Operations
Support Center ............................................................ 8-8Launch Information
Exchange Facility .......................................................... 8-10
Spaceflight Tracking & DataNetwork ........................................................................ 8-12
INTRODUCTION.
The basic purpose of mission control monitoring is to provide
guidance for launch and flight operations. This is intended to ensure
the successful achievement of mission objectives. Successful mission
accomplishment requires that launch and flight monitoring occur
during real time so that rapid decisions can be made. Data must
be analyzed rapidly in order to identify malfunctions and devia-
tions. Effects must be correlated with causes accurately, so that
corrective actions to the launch and flight operations can be taken.
Mission rules are finalized prior to the final countdown. These
rules are designed to minimize the real-time rationalizations re-
quired to cope with non-nominal situations.
MISSION CONTROL MANAGEMENT.
The mission director (MD) has overall authority during the mission.
During mission operations, the MD is usually located at the mission
control center in Houston (MCC), but he may be located at the
launch control center in Kennedy Space Center (LCC-KSC). Both
the flight and launch directors are responsible to the MD. The
. flight director at MCC exercises control over flight operationsfrom liftoff of the space vehicle through spacecraft recovery. The
launch director at LCC-KSC exercises prime control over launch
operations from the beginning of the final launch countdown until
"the space vehicle has cleared the umbilical tower. Centralizedmission control is accomplished at MCC and is supported by the
network remote stations of the spaceflight tracking and data
network (STDN), the Huntsville operations support center
(HOSC) at Marshall Space Flight Center (MSFC), and the LCC-
KSC. Specialists in the areas of navigation, electrical networks,
instrumentation, and propulsion are located at the LCC and HOSC
to provide MCC timely information as required.
MISSION CONTROL CENTER.
The flight operations director (FOD) is responsible for all opera-
tional aspects of Skylab spaceflight missions and provides the
management interface between the flight director and program
management. Direct responsibility for mission control, however,
will be vested in the flight director at the mission control center
(MCC) throughout the mission. Flight control functions will be
effected by flight controllers at the MCC utilizing the remote sites
of the STDN. Mission control will be accomplished by providing
in-flight analysis of the mission (mission trajectory, vehicle systems,
experiment systems, scientific data, and flight plan) and by con-
trolling the progress of the in-flight phase of the mission throughthe utilization of voice, telemetry, tracking, and update capabilitiesremoted from the MCC and STDN facilities.
The general mission control functions to be accomplished by flightcontrollers are as follows:
a. Monitor and evaluate, in real and delayed time, the vehicle
systems, experiment systems, scientific data, and trajectory data.
Based upon these data, decisions will be made concerning the
progress of the mission toward satisfying primary mission objectives
and mandatory, principal, and secondary detailed test objectives
(DTO's) and the need for proceeding to alternate flight plans,
contingency plans, or mission aborts.
b. Monitor and evaluate the condition of the flight crew. Based
upon these data, decisions affecting crew health and safety willbe made.
c. Perform ephemeris and maneuver updating in real time. Updated
information will be passed to the spacecraft via the up-data linkand voice from the MCC.
d. Monitor, evaluate, and update flight plan activity, including
experimental tasks, work/rest cycles, and equipment checks. Deci-
sions to alter flight plan activity will be based upon such factors
as crew status, spacecraft status, experiment status, STDN status,and weather.
e. Advise the flight crew of updated mission instructions, anomalies
in spacecraft systems found during ground monitoring, groundevaluation and recommendations to solve or circumvent any space-
craft anomalies, and recovery-area weather conditions.
MISSION OPERATIONS
CONTROL ROOM (MOCR).
MCC operations are conducted from a central MOCR where flight
controllers monitor and analyze mission status in order to makedecisions and take corrective actions consistent with flight plan
objectives and mission rules. (Figure 8-1 shows the MOCR console
layout.) Here the flight control team has rapid access to information
concerning mission progress, but are not distracted by supportfunction activities or by routine evaluation tasks. Operations in
the MOCR are assisted by several staff support rooms (SSR),
science rooms, and other supporting areas similar to those in
Apollo operations. The MOCR staff is divided into three groups
(figure 8-2): vehicle systems, experiments, and biomed; missioncommand and control; and flight dynamics/booster/
EREP/EVA. Launch vehicle responsibilities lie with three booster
systems engineers. Booster systems engineer No. 1 is responsible
to the flight director for integrating all launch vehicle activities,
executing command action as required and monitoring stage func-
tions and propulsion performance. Booster systems engineer No.
2 is responsible for integrating all stage systems. Booster systems
engineer No. 3 is responsible for integrating all IU, electrical, and
instrumentation systems monitoring and troubleshooting.
FLIGHT OPERATIONS
MANAGEMENT SUPPORT.
In the Apollo program, the spacecraft analysis (SPAN) activity
served as an interface between the flight operations organization
and the Apollo spacecraft program office (ASPO). This interface
was required by the FOD for detailed technical support from the
design, checkout and testing organizations through the ASPO rep-
resentative. This support consisted primarily of engineering judge-ment of the design personnel on the operation of the vehicle
8-1
i
Section VIii Mission Control Monitoring
MOCR CONSOLELAYOUt
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fFigure 8-1
systems in off-nominal situations. The SPAN interface also provided
the program office and other program elements with a system
for making recommendations to the flight operation and served
as a channel for receiving and evaluating certain near-real-time
information and summary data as the operation progressed. For
the Skylab program, the MSC Skylab program office will provide
a similar interface capability for the MSC spacecraft and experi-
ment hardware and for MSC flight software. Similarly, the Marshall
Space Flight Center will provide an interface for the MSFC-
designed or MSFC-contracted flight systems software, and experi-
menu. For Skylab, the designation of this function will be changed
to flight operations management support, and the designation
for the present SPAN room will be changed to the flight operations
management room (FOMR).
An additional role of the FOMR will be to accommodate the
complexities of specifying mission activity priorities (on requestfrom the flight operations team) where no clear-cut guidelines have
been established or where mission problems necessitate tradeoffs.
Thus, the FOMR operation will accomplish the following major
functions. First, it will provide detailed technical support to the
flight operations team for all flight systems hardware and software.
Second, it will have a modified role in providing required policy
level adjustments to the experiment priorities and planned mission
activities as the flight progresses. Finally, the FOMR operation
will provide support to the Skylab flight management team. The
flight operations management support organization is illustrated
in figure 8-3.
MISSION CONTROL PHASES.
Mission control during the Skylab mission phases involving the
Saturn IB vehicle are discussed below.
C.H
Prelaunch Tests.
For SL-2, two MCC command interface tests will be conducted
with each vehicle (CSM and LV). As with the SWS, this will be
the first test in which the MCC will directly interface with each
SL-2 vehicle for MCC command and telemetry verification. A
second test will be repeated as close to launch as possible to serve
as a final interface verification and will include all late telemetry
and command changes to both flight and ground hardware andsoftware. The MCC interface tests are required only for the first
CSM mission (SL-2) unless telemetry or command changes to either
onboard or ground equipment invalidate the test.
For each CSM flight the MCC may support an FRT and will
support a CDDT test.
Countdown.
The launch countdown may be monitored (voice and telemetry)
on a limited basis as early as T-24 hours. However, full MCC
support will not be provided until approximately T-6 hours.
Spacecraft systems checkout will be accomplished by KSC usingthe automatic checkout equipment (ACE), and the status and gross
system appraisal will be relayed to the MSFC and the MCC. This
information, coupled with MCC's spacecraft checkout information
and network evaluation, will provide the basis for mission GO/NO-GO decision. Specific flight control procedures during the
prelaunch phases are as follows:
a. Monitor spacecraft checkout-KSC and MCC.
b. Monitor launch vehicle systems status-launch control center
(LCC), MSFC, and MCC.
c. Verify network status-MCC and Goddard Space Flight Center
(GSFC).
8-2
Section VIII Mission Control Monitoring
MOCR ORGANIZATIONAL STRUCTURE
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Figure
d. LCC and MCC control-display status.
e. Recovery-force status and weather decision.
f. Transmit SL-I targeting parameters to the SL-2 launch vehicle
digital computer in the IU. Command loads will be generated at
the MCC based on latest tracking data and will be transmitted
to KSC at approximately T-8 hours for loading into the IU com-puter via the DCS. This load is also sent to the Huntsville opera-
tions support center (HOSC) for verification. A final update on
the target load and lift-off time will made between T-37 and T-15
minutes. This update will again be based on MCC computationof the targeting parameters from the latest available remote site
tracking data.
Both a prime and backup mode will be available for entering and
updating the targeting parameters. The prime mode is the DCS,
and the backup will be hardline via the RCA 110A computer at
KSC using a teletype (TTY) link from the MCC.
Launch.
Flight control procedures for the manned launch phase will consist
of the evaluation of spacecraft systems, launch vehicle systems,
flight crew condition, and space vehicle dynamics and trajectory.
Mission abort capability will exist for cases of adverse launch vehicle
and spacecraft system or trajectory conditions. Abort can be ini-
tiated automatically by the emergency detection system (from lift-
off to T+ 171 seconds) or manually by the crew. Abort request
can be initiated by the flight director on recommendations from
the CSM systems engineer, the booster systems engineer, and
the flight dynamics officer. Also, abort request can be initiated
by the flight dynamics officer and the booster systems engineer
when specific mission rules are violated. Abort request is relayed
to the flight crew by voice and/or by digital uplink command.
(The digital uplink command lights the ABORT light on the crew
ISKYCOM
C-N 20001 -
8-2
display panel.) Specific flight control procedures which are ac-
complished between lift-off and orbital insertion are as follows:
a. Monitor and evaluate launch vehicle systems.
b. Monitor and evaluate spacecraft systems.
c. Monitor and evaluate the condition of the flight crew.
d. Monitor and evaluate space vehicle dynamics and trajectory.
e. Call marks on abort modes.
f. Send abort request if mission rules are violated.
g. Provide contingency maneuvers.
h. Monitor the flight plan and recommend alternate procedures
to the flight crew in contingency situations.
i. Advise the flight crew of launch vehicle, spacecraft, and trajectorystatus.
LAUNCH CONTROL CENTER.
Master control of launch operations for the Skylab program is
provided by the launch complex 39 launch control center (LCC).
The LCC supported by the central instrumentation facility (CIF)
at KSC and by the Huntsville operations support center at MSFC
provides management control over the prelaunch checkout and
launch operations. After the vehicle clears the umbilical tower, the
LCC supports the MCC which then assumes control of the mission.
For a description of the launch facility, see Section VII.
LAUNCH TEAM.
The Skylab launch team (or test team) is a group of NASA
and contractor personnel responsible to the KSC center director
for performing all prelaunch, launch, and postlaunch activities for
8-3
Section VIII Mission Control Monitoring
an assigned mission. This includes, but is not limited to, the receipt,
preparation, buildup, modification, maintenance, operation, rework,
test, and checkout of individual and integrated space vehicle systems
and their associated ground support equipment, and also the prepa-
ration of required documentation. The purpose of the launch team
is to draw together all of the KSC resources required to performthe above.
The formal line/staff organization is the major functional organi-
zation from which the launch team draws personnel required to
accomplish a specific mission. The directorate of launch opera-
tions, including the launch vehicle operations and spacecraft
operations directorates, the directorate of technical support,
including the information systems and support operations direc-
torates, and supported by the directorate of installation supportare the formal organizational elements which provide the launch
team personnel. The launch team is a semi-formal organization
that has responsibility for accomplishing the operations, tests, andcountdown required for a specific mission.
The launch team operates in two modes, on-station and off-station
operations. In the on-station operational mode, the launch team
(figure 8-4) mans the test/launch consoles for the purpose of check-
ing out individual or integrated space vehicle systems and associated
ground support equipment or conducting integrated test and check-
out procedures, such as the launch countdown which requires
LCC firing room and/or ACE control room support. In the
off-station operational mode, the launch team operates duringthe period of time when the flight hardware and related support
equipment are not undergoing integrated testing, specifically during
the preparation, buildup, modification, troubleshooting; rework of
vehicle stages, spacecraft, other flight hardware, and support equip-ment.
Launch operations management is divided into two phases, test
operations planning and test operations execution. The test opera-
tions planning phase includes the responsibility of developing,
integrating, and coordinating the planning, scheduling, and tech-
nical documentation required to prepare and launch space vehicles.
The test operations execution phase includes performing the test,
checkout, and launch of the space vehicle in accordance with test
and checkout plans and procedures prepared in the planning phase.
CENTRAL INSTRUMENTATION FACILITY (CIF).
The CIF houses the following: centralized.KSC data systems for
telemetry and launch area data reception, processing, display, dis-tribution, and transmission; K SC central timing station; instrumen-
tation and standards laboratories; automatic data processing (ADP)(administrative computing) systems; and the NASA/KSC data
office. The functions performed at the CIF are a part of the KSC
MOCR
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Section viii Mission Control Monitoring
support for space vehicle testing, countdown, launch, and flight.
Prime critical power for selected electronic equipment is suppliedby the on-site generating plant.
Real Time Data Input to ClF.
The following real time data inputs are provided to the CIF tale-
merry station via KSC cable and wideband systems:
a. VHF and S-band telemetry are provided from two receivingsites. The CIF antenna site is located a mile north of the KSC
industrial area in an RF quiet zone. This site has VHF and S-band
•.ntennas, predetection recorders, receivers, and video transmission
_,f the received signals.
b. RF and hardline launch vehicle data are acquired by the launch
vehicle digital data acquisition system (DDAS) when the vehicle
is on the ground. The DDAS data are provided to the CIF telemetrystation until T-0.
c. Facility measurements from the information systems measuring
system at launch complex 39 are provided to the CIF telemetrystation.
d. Launch trajectory data from the AFETR are routed to the CIF
telemetry station. The AFETR provides vehicle position and ve-locity data.
e. The flight crew training building provides inputs from themission simulators for transmission to the mission control center
(MCC) in Houston.
f. Space vehicle orbital data are provided from MSFC via LIEF.
g. Other inputs as required to support Skylab data requirements.
Skylab Launch Data System (SLDS).
The SLDS is the information link that interconnects KSC with
MCC and other centers. The four subsystems comprising the SLDSare:
a. Countdown and.status transmitting subsystem (CASTS).
b. Television subsystem.
c. Launch trajectory subsystem.
d. Command subsystem.
The CASTS transmits up to 120 selected discrete functions and
three independent countdown time words from KSC to MCC with
an output bit rate of 2.4 kbps. These data are transmitted to MCC
independently of all other operational equipment and data circuits.
The television subsystem contains video control consoles in the
O&C building that receive 12 television channels from the launch
complex 39 operational television system. At the consoles, fivechannels are selected for transmission to the MCC. SLDS transmits
one of these channels to MCC and receives one channel from MCC.
The launch trajectory subsystem collects launch trajectory data
from AFETR and transmits it to the real time computer complex.
This subsystem provides real-time, smoothed and raw, radar tra-
jectory data from liftoff through the sub-orbital flight phase of the
vehicle. It also provides the liftoff event in the computer ID word
from the impact prediction computers at AFETR.
The command subsystem processes and transmits commands fromMCC to KSC.
LAUNCH PARAMETERS.
The criticality of the parameters monitored by the LCC at KSCfall into two major groups, mandatory and highly desirable. When
time permits, the failure of a mandatory or highly desirable item
is reported to the mission director by the launch director or the
flight director. The initial report includes the position or facility
that detected the malfunction. Subsequently, the mission director
8-6
is informed of the estimated time to repair and the .recommended
proceed, hold, recycle, or scrub action as it develops.
Mandatory.
A mandatory item is a space vehicle element or operational support
element that is essential for accomplishment of the primary mission.
These essential elements include prelaunch, flight, and recovery
operations that ensure crew safety and effective operational control,
as well as the attainment of the primary mission objectives. The
time applicable for all mandatory items is from start of countdown
to T-5 sec. If a mandatory item fails during the countdown, it
is corrected during the prelaunch phase, holding or recycling the
countdown as required. If the item cannot be corrected to permitliftoff within the launch window, then the mission director will
scrub the launch. Appropriate coordination with the launch direc-
tor and flight director, and the DOD manager for manned space
flight support operations, occurs prior to scrubbing the launch.
Highly Desirable.
A highly desirable item is a space vehicle element or operational
support element that supports and enhances the accomplishment
of the primary mission or is essential for the accomplishment of
the secondary mission objectives. The time applicable for all highly
desirable items is start of countdown to start of automatic sequence.
Consideration is given to the repair of any highly desirable item
but in no case is the launch scrubbed for any single highly desirable
item. If two or more highly desirable items fail and if other aggra-
vating circumstances occur, the mission director may scrub the
mission, following appropriate coordination with the launch and
flight directors and the DOD manager for manned space flightsupport operations.
Launch Vehicle Measurements.
The specific launch vehicle measurements that must be monitored
by KSC during the applicable time periods are listed in the "Launch
Vehicle Operations" section of the Apollo/Saturn IB Launch Mis-
sion Rules. The measurements are marked mandatory or highlydesirable to denote the type of action taken when the measurement
values go outside limits. Some inflight measurements are included
in the measurement list in addition to the hardwire redline mea-
surements. The following notes apply to the subject measurementlist:
a. Instrumentation failure is a probable cause of failure for all
parameters.
b. Minimum values are acceptable only when vehicle leakage iswithin allowable limits.
c. Minimum and maximum values represent limits of acceptableoperation.
d. The "time applicable" is the suggested time for checking aparticular redline value. Time applicable and parameter values are
chosen to simultaneously minimize countdown impact and provide
confidence that the parameters will remain acceptable to liftoff.
It is assumed that observation is continued as required.
The launch site and MCC verify, whenever possible, telemetryreadout discrepancies occurring prior to liftoff. If the MCC loses
a parameter but the launch site has a valid readout, the MCC
will continue monitoring on the basis of launch site readout. This
is true except for those mandatory parameters upon which com-mand action is taken. The measurements transmitted to MCC are
marked with an asterisk in the subject measurement list.
GSE and ESE Parameters.
A list of GSE and ESE parameters are also shown in theApollo/
Saturn IB Launch Mission Rules. A description of the malfunc-
tion/condition on parameter value limits is shown in conjunctionwith its relative times, and action to be taken is shown for eachitem.
Section Vltl Mission Control Monitoring
PRE-INSERTIONLAUNCH ABORT AREA._
50 N
40 N
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Figure 8-5
The ground support equipment (GSE) and the electrical support
equipment (ESE) included in the list are:
a. Fuel, Lox, LI-_ systems
b. Spacecraft piping
c. Pneumatic systems
d. Apollo access arm
e. Electrical power system
f. Environmental control system
g. Firing accessories
h. Umbilical tower elevator
ABORT GROUND RULES.
INTRODUCTION.Flight crew safety always takes precedence over accomplishment
of mission objectives. Abort is defined as mission termination prior
to orbital insertion as a result of a spacecraft or launch vehiclemalfunction. Unscheduled mission termination at or after orbital
insertion is referred to as early mission termination. These abort
ground rules concern primarily decisions about manual abort exe-
cution rather than automatic abort. The abort request commandis defined as a transmission from the STDN or LCC that illuminates
the red ABORT light on the command pilot's panel. This light
is one of two cues necessary for the crew to abort. A voice report
from ground control can provide the second cue. If possible, abort
request commands are based on two independent indications of
the failure. Failures are categorized according to the following
priorities:
a. Priority I failure causes destruction ofthe space vehicle, abnormal
damage to the launch complex, loss of human life, or a hazardto the astronauts.
C.fl 20004
b. Priority II failure results in damage to the space vehicle, launch
complex, or both; or requires rescheduling of the launch date.
c. Priority III failure causes a launch delay of less than 24 hours.
d. Priority IV failure has no significant effect on the launch opera-tion.
If time permits, aborts and early mission terminations are timedfor a water landing. The launch abort area extends from the launch
site to 3500 NM down range. A continuous recovery area and a
discrete recovery area are provided across the Atlantic Ocean for
recovery. Figure 8-5 shows the pre-insertion launch abort area.
AUTHORITY.Prior to liftoff, the launch director is responsible for all actions
in the event of launch site emergencies including an abort request.
The transfer of control from the launch director to the flightdirector occurs when the vehicle reaches sufficient altitude to clear
the top of the umbilical tower. Within their respective areas of
responsibility, the launch director, flight director, DOD manager
for manned space flight support operations, and mission director
may choose to take any action required for the optimum conduct
of the mission. The command pilot may initiate such action as
he deems essential for crew safety.
Hold Authority.
The command pilot, the spacecraft test conductor, the launch vehi-
cle test conductor, the launch operations manager, the launch
director, the flight director, the DOD manager for manned space
flight support operations, or the mission director may call a holdfor conditions within their respective areas of responsibility. Onlythe mission director can scrub a mission.
Abort Request Authority.
At the LCC, the launch operations manager may send an abort
request from the time the launch escape system is armed untilthe vehicle clears the umbilical tower. After tower clearance, control
8-7
Section VIII Mission Control Monitoring
shifts to MCC where the flight director, the flight dynamics officer,
or the booster systems engineer may send an abort request.
CRITERIA.
Launch Control Center.
Prior to transfer of control to the flight director, the launch
operations manager will initiate an abort request as a result ofthe following events:
a. Major structural failure or explosion
b. Loss of positive vertical motion
c. Uncontrolled vehicle tilting
d. Tower collision resulting in damage requiring immediate abort
If a vehicle collision with the umbilical tower does not require
immediate action, the launch operations manager will continue to
evaluate the damage and provide information to the flight director.
The launch operations manager will inform the flight director when
the vehicle has cleared the tower (not including the lightning rod)
by stating "clear tower" over the flight directors primary loop.
Before the launch operations manager initiates an abort request,
the following conditions must exist:
MSFCLAUNCHVEHICLESUPPORTORGANIZATION
CAPE KENNEDY
LCC
OPERATIONS
ORGANIZATION
REPRESE TArVE[r i1CIF
I[ HONTSV,LLEI i HOSCH ---- -_.i L, F LJi
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8-8Figure 8-6
a. The safety of the flight crew must be endangered.
b. An impending catastrophic condition must be observed and
reported by a periscope observer and must be confirmed by another
periscope observer or by the launch operations manager viatelevision.
c. All launch team personnel have evacuated the pad area.
d. The space vehicle has not cleared the umbilical tower.
After the abort request has been initiated, the launch operations
manager will verbally confirm to the flight crew that an abort
is recommended and will advise launch site recovery forces thatabort has been recommended.
Mission Control Center.
The flight dynamics officer will initiate the abort request for
spacecraft systems malfunctions and, when time permits, for tra-
jectory deviations and launch vehicle malfunctions. The flight
dynamics officer will initiate the abort request during the flight
phase if the vehicle exceeds the flight dynamics envelope, or if
booster cutoff conditions preclude a contingency insertion with the
SPS and time will not permit forwarding a verbal request for
initiation of this command to the flight director. The criterion
for defining the flight dynamics envelope is safe recovery of thespacecraft. The booster systems engineer will initiate the abort
request if launch vehicle malfunctions will not allow a safe insertion
and time will not permit forwarding a verbal request for initiation
of this command to the flight director.
RANGE SAFETY.
The range safety officer (RSO) can shut down the launch vehicle
engines by transmitting the manual cutoff command, which also
illuminates the ABORT light in the spacecraft. The manual cutoff
will initiate art automatic abort if transmitted prior to EDS disable.
The manual cutoff command starts a 4-see timer on the ground
that enables the destruct command capability. See Section I for
details about this system. The RSO will always safe the S-IVB
secure range safety command system upon verification of cutoff
(following a manual cutoff command) if the destruct commandis not to be transmitted.
HUNTSVILLE OPERATIONS
SUPPORT CENTER (HOSC).
During pre-mission simulations, prelaunch tests, launch countdown,
and flight of the Saturn IB vehicle, MSFC provides real-time
support through the HOSC (figure 8-6). The HOSC is arranged
on two floors in the MSFC computation laboratory as shown infigure 8-7.
Consoles in the operations support room (figure 8-8) provide real-
time vehicle status to systems engineers by means of discrete indica-
tors, analog meters, strip chart recorders, and TV displays. In
addition, all areas of the HOSC are served by voice communication
and timing systems. Consoles are permanently assigned to Skylab
orbital assembly personnel during all mission phases. Console
areas assigned to launch vehicle personnel during prelaunch tests
and launch activation will be assigned as required to other personnel
(e.g. experiment monitoring) during other mission phases. The main
conference room will house the main body of support engineersduring launch.
GROUND RULES FOR HOSC DATA DISPLAYS.
Saturn launch vehicle measurements specified to be displayed
by HOSC determine the content of the real-time data stream. In
selecting these measurements, the following ground rules are used:
a. All redline measurements in the Launch Mission Rules Document
will be displayed on the HOSC engineering console discrete indica-
Section VIII Mission Control Monitoring
HUNTSVILLE OPERATIONS SUPPORT CENTElt
HOSC
IST FLOOR
DATA J DATA DATA DISTRIBUTION
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_NAOEMEN,\_OM R_M__lk_____"_OM• ,'_ --m-Ai
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DURING SUSTAINED ORBIT
HOSC
2ND FLOOR
Figure 8-7
SUSTAINED ORBIT
Section VIII Mission Control Monitoring
OA ASTN LV
AUX EC S/tCS/TAC S SII/StV8 SiC/StB
mB
ASTR OA ASTR LV
Figure 8-8
/
1C-N 20006
tors continuously from the start of the terminal countdown untillaunch. Each redline discrete indicator will illuminate if the real-
time telemetry data indicates that the measurement has violated
the minimum or maximum value specified in the Launch Mission
Rules Document. Redline indicator displays are identified by alabel with a solid red colored background.
b. All launch vehicle measurements identified as MANDATORY
or HIGHLY DESIRABLE inflight measurements in the Launch
Mission Rules Document will be displayed on the HOSC engi-
neering console display devices in order to provide real-time supportto KSC and MCC.
c. Flight control measurements will be displayed on the HOSC
engineering console display devices (discrete indicators, analog
meters, strip-chart recorders, and/or digital/TV formats).
d. Measurements that are required to support the Saturn launch
vehicle prelaunch testing, countdown, and flight operations, and
measurements that provide mission sequence of events (commands,
events, and system response) and/or provide a basis for real-time
launch vehicle systems performance evaluation will be displayedin the HOSC.
e. Measurement displays are limited by: (1) the availability ofconsole display devices and (2) the capacity of the real-time data
frame (1023 10-bit words).
CONSOLE SUPPORT ENGINEERS.
Support is provided as required to launch and flight operations
from HOSC engineering data consoles. The consoles are manned
by systems engineers identified by MSFC Laboratories, who per-
form detailed system data monitoring and analysis and coordinate
support to the Saturn launch vehicle staff support room in
mission control center and engineer-to-engineer support to KSC.An observer from each launch vehicle stage contractor is alsolocated at the consoles.
SYSTEMS SUPPORT ENGINEERS.
Systems support engineers are organized into preselected subsys-
tem problem groups including technical and management personnel
who provide support in launch vehicle areas that may be the subject
of a KSC or MSC request for analysis during terminal launch
countdown and flight. The System° support engineers are com-
posed of both MSFC and stage contractor personnel.
8-10
Hose COMMUNICATOR.
The HOSC communicator is a KSC launch vehicle operations
member who mans the LIEF console in the CIF during the terminal
countdown and coordinates with the HOSC all requests for supporton the launch vehicle from KSC launch vehicle operations person-
nel. He also provides launch vehicle status information during the
countdown to the launch vehicle operations communicator andcontrols switching of operational TV to HOSC.
SIMULATION SUPPORT.
A number of support activities will provide special systems simu-
lation functions to supplement MCC capabilities. These activities
generally utilize existing engineering development hardware and/ormath models not provided at the MCC and provide support on
contingency and/or scheduled basis. Those activities related to
support of the Saturn IB vehicle are as follows.
Wind Monitoring Team.
The wind. monitoring team monitors measured ground and flight
wind observations at KSC and provides GO/NO-GO recom-
mendations and advisory support to KSC in the event of marginal
conditions for launch and certain other pad activities. Vehicle
structural bending moment effects are calculated from ground
winds, and compared against vehicle capabilities. The actual flight
wind profiles, measured prelaunch, are used in a flight dynamics
math model to predict the acceptability of structural and control
effects. Predicted vehicle in-flight dynamic responses are provided
to the MCC flight controllers and the crew. Coordination with a
similar spacecraft team at MSC will be maintained for the manned
launches and joint recommendations issued.
Saturn IB Systems Development Facility.
The Saturn IB system development facility (breadboard) will be
held in a standby mode during major prelaunch test and countdown
activities to assist in troubleshooting and testing irt the event of
problems at KSC. This facility is used pre-mission to develop
and verify mission software, and include actual or breadboard
launch vehicles and ground computer hardware, software, and
related systems in the configuration similar to that utilized at KSC.
Target Update Team.
The target update team, working closely with the MCC flight
dynamics personnel, will verify the performance capability of the
Saturn IB vehicle to achieve the updated rendezvous target condi-
tions generated by the MCC from tracking of the orbital assembly.
The MCC will generate updated Saturn IB guidance targeting
parameters after launch of SL-I and transmit these to the SL-2
and subsequent vehicles at KSC shortly before their launch. The
target update team will verify the performance acceptability of
the update conditions for MCC and KSC.
OPERATIONS SUPPORT PHASES.
The level of support activities will vary significantly as the total
mission activities change. The mission can be divided into five
characteristic support phases, some of which are repeated over the
course of the mission. These phases and their occurrence during
the mission are summarized in figure 8-9.
LAUNCH INFORMATIONEXCHANGE FACILITY (LIEF)
The LIEF is a network of communications resources providing closeday-to-day exchange between KSC and MSFC of information
relating to launch vehicle integration, checkout, and successful
mission accomplishment. Direct engineering support is provided
in areas of propulsion, navigation, and electrical networks. The
joint MSC/MSFC wind monitoring team of the LIEF organization
advises the launch director of the acceptability for launch of theactual wind environment.
Section VIII Mission Control Monitoring
OPERATIONSSUPPORT PHASES
SUPPORT PHASES VEHICLE ACTIVITIES
I PREMISSION PHASE SL-I PRELAUNCHSL-2 pRELAUNCH
II SL-1 LAUNCH PHASE SL-1 LAUNCH *SL-2 PRELAUNCH
SL-2, SL-3 ANDIII SATURN IB LAUNCH SL-4 LAUNCHES *
PHASE OA UNMANNED
IV MANNED OPERATIONSPHASE OA MANNED
V ORBITAL STORAGE OA UNMANNED
PHASE SL-3 AND SL-4PRELAUNCH
• LAUNCH PERIODS DEFINED FROM LIFTOFFTHROUGH IU ACTIVE LIFETIME
• * SIMULATIONS, MCC INTERFACE TESTS
PHASE I
ACTIVE OPERATIONS SUPPORT INTERFACES
SATURN
KSC
SKYLAB
8SE FOMR
PHASE tl'_j 1_ PHASE Ill
It PHASE'V I
LAUNCH
PHASE V
DURATION
6 MONTHS
1 DAY
12-30 HOURS
SL-_ 27 DAYS
SL-3:55 DAYS
SL-_ 55 DAYS
42-47 DAYS
r.._PHASE III Ii IF/PHASE IllPHASE ,V , PHASE V PHASE IV
SL-3 SL-4 / END OFLAUNCH -- MISSION
LAUNCH
Figure
ADVISORY SUPPORT.
An advisory group is convened at the MSFC Huntsville operations
support center (HOSC) conference room during countdown and
major tests to provide timely response to KSC engineering support
requests. Primary communications are by a voice line between the
HOSC operations room and the LIEF console at launch complex3_%
REAL-TIME DIGITAL DATA.
MSFC is supplied real-time telemetered data from either of two
data cores in the CIF by means of the LIEF wideband circuit
(40.8 kbs). MSFC controls parameter selection via the LIEF real-
time data request, a serial PCM message transceived at 2.4 kbsvia 201B MODEM link.
The real-time data output frame contains 1023 data words of 10
bits each, plus a 30-bit sync word. A complete data request in-
struction assigns a parameter in each of these 1023-word slots by
designating the storage memory address for each. The complete
instruction is stored in the data request memory and executed
automatically at the 40.8 kbs rate, which is roughly four framesper second. MSFC updating can at any time change one word,
a dozen words, or all 1023 words in the data request memory.
MSFC will provide raw space vehicle orbital data from the MSFN
to KSC over existing data transmission circuits (MSFC to KSC)
at 50 kbs via 303G MODEM link. The STDN data will be displayedat KSC in real-time.
TAPE-TO-TAPE INFORMATION EXCHANGE.
The KSC branch of the information exchange system consists of
two Univac 1005 computers at the CIF, interconnected by the 301BMODEM links. Similar installations at MSFC interconnect with
the KSC system to form a versatile high-speed (40.8 kbs) digitaldata information exchange system.
C-H 2001
8-9
IBM 066-068 CARD TRANSCEIVER NETWORK.
The card-to-card transceiver transmits over a WATS line (4kHz)
to any station equipped with a matching unit. The KSC transceiver
is in the CIF. The data rate is 10 to 12 cards per minute.
FACSIMILE NETWORKS.
NASA leases alternate-voice facsimile equipment of two types for
transmission and reception of documents, drawings, and continuous
tone images over standard 4 kHz voice circuits. Magnafax Model
850 stations, each capable of transmitting or receiving approxi-mately 10 pages (8-1/2" x 11") per hour, make up one network.
A station directory is published by the NASA network contractor,
and calls are made in much the same manner as placing a telephonecall.
A high-speed LDX facsimile network links NASA headquarters,
KSC, MSC, MSFC, and various NASA contractors through switch-
ing facilities located at MSFC. Duplex terminal equipment at each
location permits simultaneous transmission and reception of copy
at the rate of approximately 60 pages (8-1/2" x 11") each wayper hour.
COUNTDOWN TIMING AND LIFTOFF.
KSC countdown timing from the LCC timing racks is transmittedto countdown clocks in the MSFC HOSC via the LIEF console.
The LIEF voice circuit is allocated to this 1700 Hz, serial-PDM,
encoded signal. Another circuit provides MSFC with a liftoffsignal.
CLASSIFIED TELETYPE.
An on-line teletype facility interconnects the KSC communications
center in the KSC headquarters building with NASA headquarters
and other NASA centers. This is the only facility for exchange
8-11
Section VIII Mission Control Monitoring
of classified launch operations information between NASA ele-ments.
CLOSED CIRCUIT TV.
The LIEF console operator selects and monitors the TV imagerequested for countdown and launch coverage by the HOSC com-municator.
LIEF CIRCUIT RECORDING.
Certain voice circuits are recorded during the countdown for sub-
sequent analysis of launch operations. The LIEF video circuit is
also recorded during launch.
SPACEFLIGHT TRACKING &DATA NETWORK (STDN).
Tile STDN, managed by Goddard Space Flight Center, will be
utilized during the Skylab missions. This network consists of 13
land sites (figure 8-10) and the ship Vanguard. These sites are
situated around the earth to provide as nearly continuous com-
munication with the vehicles as possible.
The basic types of support required from the network are tracking,
telemetry record, telemetry real-time display, command communi-
cation, voice communication, and television. Figure 8-I 1 provides
a matrix showing station capabilities.
SPACEFLIGHTTRACKING AND DATA NETWORK
60o
30o •
0 o
30o
180°
I
150° 120 °
LAUNCH AREA
CIP - CENTRAL INSTR FACILITY
MLA - AFETR C-BAND
CNV - CAPE KENNEDY
MIL- MERRIT ISLAND
ACN - ASCENSION ISLAND
BDA - BERMUDA
90° 60 ° 30o 0 o 300 600 90° ! 20 ° 150° 180 °
5 CYI - GRAND CANARY ISLAND 10 MAD - MADRID, SPAIN
6 CRO - CARNARVON, AUSTRAUA 11 GWN - GUAM
7 GDS - GOLDSTONE, CALIF. 12 TEX - CORPUS CHRISTI, TEX.
8 HAW - HAWAII 13 NFL - NEWFOUNDLAND
9 HSK - HONEYSUCKLE, AUSTRALIA
Figure 8-10
C-H 14488-1
8-12
Section VIII Mission Control Monitoring
STDN STATION CAPABILITIES
TRACKING
C-BAND RADAR
USB/CCS
TELEMETRY
VHF
USB/CCS
DATA PROCESSOR
DATA REMOTI NG
RIOMED REMOTING
DATA RECORDI NG
COMMAND
UHF
USR/CCS UPDATA
CMD PROCESSOR
CMD REMOTING
A/G VOICE
VHF
USR
TV
S-BAND TV RECORD
R/[ TRANSMISSION
CNV CIF MIL MLA BDA CYI ACN MAD CRO GWM HSK HAW GDS TEX VAN NFL
Figure 8-11
CA 200Q8
8-13
TABLE OF CONTENTS
Flight Profile ....................................................................
Hold & Recycle Rullngs ................................................Launch Wlndow ..............................................................
Environment ....................................................................
Fllght Mlsslon Rules ......................................................EDS Llmlts Derlvatlon ....................................................
S-IVBIIU Orbltal Decay ..................................................
S-IU Active Lifetime ........................................................S-IVB Active Lifetime ....................................................
Orbital Configuration ......................................................
k. C-Band track of the S-IVB/1U will be required every TBD
revolutions from end of IU life until reentry. This support will
9-1 be supplied on a non-interference basis with SWS skin tracking
9-1 support.
9-1 1. The S-IVB auxiliary propulsion system will provide attitude
9-1 control capability for 7 hr 30 min for SL-2 and SL-4 and 4 hr9-4 30 min for SL-3.9-4
FLIGHT PROFILE.
9-5
9-5
9-5
9-5
Several constraining factors govern the Saturn IB launch vehicle
flight profile to achieve the desired orbit insertion conditions such
as vehicle performance, targeting conditions, range support, and
launch complex design. A summary of these factors follows:
a. The SL-2 launch will occur from launch complex 39B into a
50-deg inclination orbit. Since Pad 39B is built with the launchazimuth fixed at 90 deg, a roll command will be required to provide
the flight azimuth required (see launch window discussion).
b. To provide proper initial conditions for phasing orbit, the S-IVB
cut-off conditions will be targeted for an 81 x 120 NM orbit inser-
tion. The nominal insertion orbit must have a perigee greater than70 NM.
c. The launch vehicle or spacecraft will not be allowed to enter
a flight regime, as defined by the Flight Mission Rules, from which
a safe abort cannot be accomplished.
d. The nominal ascent trajectory will be shaped so that aerody-
namic loads during a full-lift, free-fall abort from any point along
the trajectory shall not exceed 16 G's.
_. The trajectory will be shaped to maintain the aerodynamic
heating indicator below an integrated value of 63.6 x 10 7 N-m/m -"
(43.6 x 10 6 lbf-ft/ft.:).
f. For launch aborts, the duration of free fall above the entry
interface of 300,00 ft will be at least 100 sec to provide adequate
time to perform functions necessary for safe spacecraft entry.
g. To provide compatibility between guidance commands and ve-
hicle reaction, the guidance command angles will be rate limited
to 1 deg/sec during powered flight.
h. Continuous range safety command destruct and sating capability
is required from launch minus 30 min through insertion plus 60
sec.
i. Continuous STDN coverage of the launc h vehicle is desired from
start of automatic countdown until 60 sec after insertion.
j. One 3-min pass for telemetry, command, and tracking coverageis desired for the S-IVB stage of the Saturn IB during each revolu-
tion through IU powered operation.
HOLD & RECYCLE RULINGS.
Unscheduled hold and recycle rulings depend not only on the type
of malfunction, but also on the time period in which the malfunctionoccurs.
The terminal launch countdown will he held or scrubbed to repair
any loss of redundancy if an inflight failure of the affected system,
subsystem or function could:
(a) Compromise crew safety, or
(b) Require early termination of the mission, or
(c) Require cancellation or early termination of subsequent mis-sions.
Figure 9-1 lists the actions taken for particular malfunctions occur-
ring at various times from start of space vehicle countdown untilliftoff.
LAUNCH WINDOW.
The Saturn IB launch windows are constrained by the phase anglebetween the SWS and the CSM at insertion and the amount of
Saturn IB payload capability sacrificed for the propellant used for
yaw steering. Length of the SL-2 launch window is based upon
the SL-1 descending node being 153.25 deg, a 5 to 7 orbit rendez-
vous scheme, and a yaw steering allowance of 700 Ibm of propellant.
The 700 Ibm of yaw steering propellants represent approximately
16 min of launch window capability. The flight azimuth range for
this launch window is between 51.82 deg at the opening of the
window and 37.68 deg at the close of the window. The launch
window times shown in figure 9-2 are current approximations.
ENVIRONMENT.
The accomplishment of mission objectives are contingent upon
appropriate environmental conditions. Surface wind, upper air, and
weather contingencies are considered below.
SURFACE WIND RESTRICTIONS.
Preliminary results from wind tunnel tests conducted on a 5.5%
scale aeroelastic model of the SL-2 vehicle and pedestal in the
9-1
Section IX Mission Variables and Constraints
TIME PERIOD
I. T-26 HR (START OF S/_/COUNT)
TO T-12 HR 15 MIN (START OF
LV COUNT).
2. T-12 HR 15 MIN (START OF LV
COUNT) TO T-6 HR 30 MIN.
5.
T-6 HR 30 MIN (START S-IB
LOX LOADING) TO T-4 HR
15 MIN (START S-IVB LH 2 LOAD-ING).
T-4 HR 15 MIN (START OF LH 2
LOADING) TO T-3 HR 15 MIN
(SC, START OF CABIN CLOSE-
OUT); TO I"-30 MIN (LV).
T-3 HR 15 MIN (START OF SC
CABIN CLOSEOUT) TO T-30
MI N (CLEAR ACCESS ARM).
6. T-30 MIN (CLEAR ACCESS ARM)
TO T-10 MI N (S-IVB THRUST
CHAMBER CHI LLDOWN).
7. T-10 MIN (S-IVB) THRUST
CHAMFER CHILLDOWN TO T-3
SEC (IGNITION).
FUNCTION/CONDITION
MALFUNCTION
MALFUNCTION OF ANY REPAIRABLE SPACECRAFT
SYSTEM.
MALFUNCTION OF ANY REPAIRABLE SPACE
VE HICLE SYSTEM.
MALFUNCTION OF ANY REPAIRABLE SPACE
VEHICLE SYSTEM.
MALFUNCTION OF ANY REPAIRABLE SPACE
VEHICLE SYSTEM.
PROBLEM IN SPACECRAFT CABIN CLOSEOUT.
MALFUNCTION OF ANY REPAIRABLE SPACE
VEHICLE SYSTEM.
ANY MALFUNCTION OF THE S/_/ GUIDANCEAND CONTROL SYSTEM.
ANY MALFUNCTION OF THE EDS.
EXCEEDING OR LOSS OF ANY SPACE VEHICLE RED-
LINE VALUE APPLICABLE TO THIS TIME PERIOD.
A. LOSS OF ANY COMPLETE S/V TELEMETRY
LINK TO T-2 MIN 43 SEC.
B. LOSS OF ONE OR MORE OF THE FOLLOW-
ING TELEMETRY LINKS TO T-3 SEC, S-IB LINK
GP-1, S-IVB LINK CP-1, IU LINK DP-1, ANY
COMPLETE S/C LINK.
LOSS OF ANY MANDATORY MEASUREMENT.
MALFUNCTION OF VEHICLE FIRE DETECTION SYSTEM.
(2 OF 4M, 3 OF 4HD).
MALFUNCTION OF CIF TM STATION SUCH THAT THE
GO/NO-GO STATUS OF THE LAUNCH VEHICLE TM
SYSTEMS CANNOT BE DETERMINED.
NON-OPERATION OR MALFUNCTION OF THE DIGITAL
DATA ACQUISITION SYSTEM WHICH COULD RESULT
IN THE PRESENTATION OF ERRONEOUS DATA TO THE
OPERATIONAL COMPUTERS AND/OR ESE DISPLAYDEVICES.
ACTION/COMMENT
PROCEED. CORRECT MALFUNCTION IN PARALLEL WITH OTHER
OPERATIONS. HOLD AT T-12 HR 15 MIN FOR MANDATORY OR
HIGHLY DESIRABLE ITEMS IF ESTIMATES FOR COMPLETION IN-
DICATE ALL WORK (REPAIR AND CLOSEOUT) CANNOT BE ACCOM-
PLISHED PRIOR TO T-6 HR 30 MIN.
PROCEED. CORRECT MALFUNCTION IN PARALLEL WITH OTHER
OPERATIONS. HOLD AT T-6 HR 30 MIN FOR HIGHLY DESIR-
ABLE OR MANDATORY ITEMS IF SERVICE STRUCTURE IS REQUIR-
ED FOR REPAIR. HOLD AT T-5 HR 45 MIN IF ONLY THE AC-
CESS ARM IS REQUIRED FOR REPAIR.
PROCEED OR HOLD. PROCEED IF CORRECTION CAN BE ACCOM-
PLISHED IN PARALLEL WITH NORMAL FUNCTIONS:OTHERWISE
HOLD FOR MANDATORY AND HIGHLY DESIRABLE ITEMS.DURING
THE HOLD REVIEW CRITICALITY WITH REFERENCE TO PERFORM-ANCE DEGRADATION, EVALUATE REPAIR TIME WITH RESPECT
TO LAUNCH WINDOW TO DETERMINE NECESSITY FOR SCRUB
(I HR TO RETURN SERVICE STRUCTURE).
PROCEED OR HOLD. PROCEED IF CORRECTION CAN BE ACCOM-
PLISHED IN PARALLEL WITH NORMAL FUNCTIONS: OTHERWISE
HOLD FOR MANDATORY AND HIGHLY DESIRABLE ITEMS. RE-
PAIR IF POSSIBLE WITHOUT THE USE OF SERVICE STRUCTURE. IF
REPAIR IS NOT POSSIBLE WITHOUT SERVICE STRUCTURE, REVIEWCRITICALITY: MAKE DECISION AS TO PERFORMANCE DEGRA-
DATION, SCRUB IF MANDATORY. HOLD AT T-3 HR 15 MIN FOR
COMPLETION OF INTERNAL CM WORK IF REQUIRED.
HOLD AT T-30 MIN FOR COMPLETION OF CABIN CLOSEOUT.
PROCEED OR HOLD. PROCEED IF CORRECTION CAN BE
ACCOMPLISHED IN PARALLEL WITH NORMAL FUNCTIONS AND
ACCESS TO THE PAD IS NOT REQUIRED; OTHERWISE HOLD AT
T-10 MIN FOR MANDATORY AND HIGHLY DESIRABLE ITEMS.
REPAIR _F POSSIBLE WITHOUT THE USE OF SERVICE STRUCTURE.
IF REPAIR IS NOT POSSIBLE, REVIEW CRITICALITY; MAKE DE-
CISION AS TO PERFORMANCE DEGRADATION, SCRUB IFMANDATORY.
T- 10 MI N TO T-3 MI N 7 SEC (LAUNCH SEQUENCE START):
HOLD. A HOLD OF 5MIN MAXIMUM CAN BE TOLERATED
WITHOUT RECYCLING TO T-10 MIN PROVIDED S-IVB CHILL-
DOWN OPERATIONS CONTINUE. THIS LIMITATION IS DiC-
TATED BY S-IVB THRUST CHAMBER CHILLDOWN. IF THE HOLDEXCEEDS S MIN RECYCLE TO T-10 MIN. A WARMUp PERIOD
OF 10 MIN IS REQUIRED PRIOR TO INITIATING CHILLDOWNSEQUENCE AFTER RECYCLING TO T-10 MIN.
T-3 MIN 7 SEC TO T-3 SEC:
CUTOFF.
RECYCLE TO T-10 MIN.
MAKE THE DECISION TO HOLD AND REPAIR OR TO SCRUB.
A WARMUP PERIOD OF 10 MIN IS REQUIRED BEFORE INITI-
ATING S-IVB CHILLDOWN SEQUENCE AFTER RECYCLING TO
T-10 MIN.
NOTE: FROM A TECHNICAL STANDPOINT THERE IS NO
LIMIT TO THE NUMBER OF RECYCLES TO T-10 MINUTES THAT
THE LAUNCH VEHICLE CAN WITHSTAND IN THE TIME PERIOD
T-10 MINUTES TO T-3 SEC PROVIDING RED-LINE VALUES
REMAIN WITHIN LIMITS.
Figure 9-1 (Sheet 1 of 2)
C-H 14559.1
9-2
presence of a scale model of the LUT indicate that the maximum
SL-2 vehicle response should be of no greater magnitude than that
of SA-205 for comparable wind speeds. By launching SL-2 from
LC-39, the effect of umbilical tower influence on the vehicle re-
sponse was changed from that of LC-34 (SA-205 launch), and theLC-39 wind damper can be connected to the SL-2 vehicle for
operations prior to launch. Maximum vehicle response loads occur
on SL-2 with the empty vehicle damper-off configuration, which
is comparable to the empty SA-205 configuration. Based on the
vehicle response to ground winds and the SA-206 structural charac-
teristics, the preliminary prelaunch and launch wind limits, com-
pared to SA-205, are shown in figure 9-3.
UPPER AIR RESTRICTIONS.
Preflight simulations of the space vehicle response to upper air
Section IX Mission Variables and Constraints
winds at the time of launch are performed at MSFC by an MSC-
MSFC Wind Evaluation Team using wind data provided by KSC.
The Cape FPS-16, radar, and the 1.16 (or a suitable replacement
in the event 1.16 is not operational) are used to track jimsphere
balloons. Jimsphere balloons are released from the launch pad toobtain launch wind information from T -50 hr to T + 10 min
on a schedule agreed to by KSC, MSFC, and MSC. Results of
the MSFC-MSC wind simulations are provided through the LIEF
coordinator to the launch operations manager, or in his absence
the test supervisor, in the launch control center. The results may
be forwarded periodically as considered appropriate until the re-
lease at T - 10 hr. A report is provided for each balloon releasedfrom T - 10 hr to T -5 hr 30 min. Results of the wind simulations
are summarized in writing in the HOSC and transmitted by datafaxto the LIEF coordinator in the LCC.
If simulations indicate that wind conditions are marginal for a
TIME PERIOD
8. T-10 MIN (S-_VBTHRUST
CHAMBER CHILLDOWN) TO T-3
SEC (IGNITION)
9. T-10 MIN (S-IVBTHRUST
CHAMBER CHILLDOWt_ TO
T-3 SEC (IGNITION).
10. T-3 SEC (IGNITION) TO T-0
(LIFTOFF)
FUNCTION/CONDITIONMALFUNCTION
ACTION/COMMENT
ANY MALFUNCTION OF THE S/C REACTION
CONTROL SUBSYSTEM.
ANY MALFUNCTION OF THE S/C SERVICE
PROPULSION SUBSYSTEM.
ANY MALFUNCTION OF THE S/C ELECTRICAL
POWER SUBSYSTEM.
ANY MALFUNCTION OF THE S/C ENVIRONMEN-
TAL CONTROL SYSTEM.
MALFUNCTION OF S/C C-BAND BEACON.
MALFUNCTION OF L/_/C-BAND SYSTEM OR
GLOTRAC BEACON TO T-2 MIN 43 SEC.
MALFUNCTION OF MILA USB STATION
MALFUNCTION OF S/C S-BAND SUBSYSTEM
MALFUNCTION OF LAUNCH VEHICLE RANGE
SAFETY COMMAND SYSTEM (ALL FOUR CDRS
ARE REQUIRED).
LOSS OF ETR GROUND SUPPORT INSTRUMENTA-
TION FOR THE SPACE VEHICLE COMMAND SYSTEMS
(6428 COMPUTER, DRUL, FRW-2A, TRANSMISSION
LINES).
LOSS OF COUNTDOWN CLOCK (GMT) OUTPUT.
MALFUNCTION OF FLYWHEEL GENERATOR.
A. LOSS OF GROUND COMPUTER SYSTEM
SERVICE AND DEE-6 SERVICE TO T-20 SEC.
B. LOSS OF GROUND COMPUTER SYSTEM
SERVICE OR DEE-6 SERVICE TO T-3 SEC.
C. LOSS OF AGCS COMPUTER DISCRETE OUT-
PUT POWER TO T-3 SEC.
MALFUNCTION OF THE AALT TO T-5 SEC.
EXCEEDING ANY WEATHER RESTRICTION
APPLICABLE TO THIS TIME PERIOD.
LOSS OF ANY MANDATORY COMMUNICA-
TION LINK APPLICABLE TO THIS TIME PERIOD.
NOT APPLICABLE.
T-1O MIN TO T-3 MIN 7 SEC (LAUNCH SEQUENCE START).
HOLD. A HOLD OF 5 MIN MAXIMUM CAN BE TOLERATED
WITHOUT RECYCLING TO T-10 MIN PROVIDED S-IVB CHILL-
DOWN OPERATIONS CONTINUE, THIS LIMITATION IS
DICTATED BY S-IVBTHRUST CHAMBER CHILLDOWN. IF
THE HOLD EXCEEDS S MI N. A WARMUP PERIOD OF 10 MIN
IS REQUIRED PRIOR TO INITIATING CHILLDOWN SEQUENCE
AFT ER RECYCLI NG TO T- 10 MI N.
T-3 MIN 7 SEC TO T-3 SEC.
CUTOFF.
RECYCLE TO T-10 MIN.
MAKE THE DECISION TO HOLD AND REPAIR OR TO SCRUB. A
WARMUP PERIOD OF 10 MIN IS REQUIRED BEFORE INITIATING
S-IVB CHILLDOWN SEQUENCE AFTER RECYCLING TO T-10 MIN.
NOTE: FROM A TECHNICAL STANDPOINT THERE IS NO LIMIT
TO THE NUMBER OF RECYCLES TO T- 10 MI N THAT THE LAUNCH
VEHICLE CAN WITHSTAND IN THE TIME PERIOD T-10 MIN TO
T-3 SEC PROVIDING RED-LINE VALUES REMAIN WITHIN LIMITS.
T-10 MIN TO T-3 MI N 7 SEC (LAUNCH SEQUENCE START).
HOLD. A HOLD OF 5 MIN MAXIMUM CAN BE TOLERATED
WITHOUT RECYCLING TO T-10 MIN PROVIDED S-IVB CHILL-
DOWN OPERATIONS CONTINUE. THIS LIMITATION IS
DICTATED BY S-IVB THRUST CHAMBER CHILLDOWN_ IF THE
HOLD EXCEEDS 5MIN. A WARMUP PERIOD OF 10 MIN IS
REQUIRED PRIOR TO INITIATING CHILLDOWN SEQUENCE
AFTER RECYCLING TO T-10MIN.
T-3 MIN 7 SEC TO T-3 SEC
CUTOFF.
RECYCLE TO T-10 MIN.
MAKE THE DECISION TO HOLD AND REPAIR OR TO SCRUB. A
WARMUP PERIOD OF 10 MIN IS REQUIRED BEFORE INITIATING
S-IVB CHILLDOWN SEQUENCE AFTER RECYCLING TO T-10 MI N.
NOTE: FROM A TECHNICAL STANDPOINT THERE IS NO LIMIT
TO THE NUMBER OF RECYCLES TO T-10 MIN THAT THE LAUNCH
VEHICLE CAN WITHSTAND IN THE TIME PERIOD T- 10 MIN TO
T-3 SEC PROVI DI NG RED- LINE VALUES REMAI N WIT HI N LI MIT S.
NONE.
NO HOLDS WILL BE CALLED.
NO MANUAL CUTOFF WILL BE GIVEN.
AN AUTOMATIC CUTOFF WILL RESULT IN A SCRUB.
Figure 9-1 (Sheet 2 of 2)
C-H 14563-1
9-3
Section IX Mission Variables and Constraints
OPPORTUNITY FROM TO
(HR: MI N: SEC) (HR:MIN:SEC)
PRIME 1 1 MAy 73 12 NOON 12:09:00 PM
ALTERNATE 2 1 MAY 73 11:28:30 AM 11:34:30 AM
ALTERNATE 6 6 MAY 73 1_.00 AM 10:.09 AM
ALTERNATE 7 7 MAY 73 9:29 AM 9:35 AM
NOTE: TIME REFERENCED TO EASTERN STANDARD TIME (EST).
Figure 9-2
C-H 20145
safe launch, the MSC-MSFC team includes in their report that
"Launch winds are marginal for launch" prior to T -2 hr 15 min.
Upon receipt of this message, the launch operations manager places
a contingency plan into effect that will provide for a new balloon
release every hour. The contingency plan remains in effect until
liftoff has occurred, until the launch has been scrubbed, or until
a subsequent MSC-MSFC wind report states "Launch winds are
no longer marginal for launch". The MSC-MSFC team provides
a report to the launch operations manager for each release under
the contingency plan. Balloon release information is updated prior
to the Flight Readiness Report (FRR). Figure 9-4 shows the typical
wind speed limits for Skylab Saturn IB/CSM Flights.
WEATHER RESTRICTIONS.
It is highly desirable that launch site area visibility be sufficient
for at least one of the long range ground tracking cameras to have
an unobstructed view of the space vehicle through the period of
maximum dynamic pressure. Failure of optical tracking systemsor weather conditions that would result in failure to meet this
requirement are considered basis for a hold. It is mandatory that
visibility be sufficient for the forward observers to monitor space
vehicle conditions. A minimum visibility of 3 NM and a ceilingof 500 ft is considered mandatory.
The space vehicle will not be launched if the nominal flight path
will carry the vehicle:
UNPRESSURIZED PRESSURIZED
F.S.F. 1.4 51.2 [64] (53) 51.5 [64]
1.25 51.8 [7 2] (54) 73.8 (56)
F.S.E. 1.4 33.0 [5'] (30) 72.5
1.25 33.2 [6 3] (33) 77.4
LAUNCH 29.0 (27)*
VELOCITIES EXPRESSED IN KNOTS AT 530-FT LEVEL
[ ] = DAMPER ENGAGED
( ) = AS-205 LIMITS CORRECTED TO 530-FI REF LEVEL
* = AS-20S LAUNCH LIMtT BASED ON SPACECRAFT NOT
STRUCTURAL CRITERIA
F.S.F. = FREE STANDING FULL
F.S.E. = FREE STANDING EMPTY
Figure 9-3
C.H 20146
a. Within five statute miles of a cumulonimbus (thunderstorm)cloud or within three statute miles of an associated anvil.
b. Through cold-front or squall-line clouds that extend above10,000 ft.
c. Through middle cloud layers 6,000 ft or greater in depth wherethe freeze level is in the clouds.
d. Through cumulus clouds with tops at 10,000 ft or higher.
FLIGHT MISSION RULES.
Flight Mission Rules are procedural statements which provide flight
control personnel with guidelines to expedite the decision-making
process. The rules are based on an analysis of mission equipment
configuration, systems operations and constraints, flight crew pro-
cedures, and mission objectives. The rules can be categorized asgeneral and specific. The general mission rules contain the basic
philosophies used in the development of rules while the specificmission rules provide the basic criteria from which real-time deci-
sions are made. The Cues/Notes/Comments column of specific
mission rules and the background data and flight control procedures
provide the flight controller with additional information concerning
the condition/malfunction and ruling.
EDS LIMITS DERIVATION.
The EDS limits are established to assure safe separation of the
CSM from the launch vehicle in an emergency situation. The
Aero-Astrodynamics Laboratory at MSFC determines the EDS
limits for the Saturn IB launch vehicle by math model simulationsof various failure modes. The simulations determine the vehicle
breakup rates; then, the EDS limits are selected between the three-
sigma nominal and the breakup rates. See Section Ill for EDSsettings and additional EDS information.
|1 d| | 111 i,'ll !:! i,',[o].11 :i _11 :i[i I |'JI:111 i lip] |11 I_f,;I iO" 120
_. 4o
I I I /_WiND SLOWING FROM
NORTH CORRESPONDS TO
AN,_,MUTHOF0O / \/ \
/ \
--J yR'GHTC SSWINO/X
.HEA_W, ND k _A,Li, iD L SS_IN_Di
o _-_-_ I I1 In-CRIC_I0 80 160 240 320
WIND AZIMUTH (DEG FROM NORTH)
---- WIND SPEED LIMIT
-- 99 PERCENTILE JANUARY WIND
----- 99 PERCENTILE FESRUARY WIND
..... 99 PERCENTILE MARCH WIND
Figure 9-4
C-H 20147
9-4
S-IVB/IU ORBITAL DECAY.
The nominal lifetime for the SL-2 S-IVB stage launched into an
81 by 120 NM orbit is 23.1 hr (0.96 days) or approximately 15
revolutions. The plus two-sigma and minus two-sigma densities
give lifetimes of 21.6 hr (0.90 days) and 24.8 hr (1.03 days), respect-
ively. The altitude decay history for the SL-2 S-IVB stage launched
May 1, 1973, is shown in figure 9-5.
IU ACTIVE LIFETIME.
The IU has a design-goal lifetime of 6 hr 42 min based on a 0.992
reliability figure. However, analyses of individual component and
subsystem life expectancies indicate that the pacing item is the
6D40 battery. With the expected electrical load profile, this battery
should last approximately 11 hr. Another limiting factor is the GN 2
usage rate, which should deplete the stored supply in approximately14 hr.
Section IX Mission Variables and Constraints
sec after J-2 engine cutoff) the vehicle consists of the SIVB stage,
IU, SLA, and CSM. The S-IB stage separated, the ullage rockets
jettisoned, and the launch escape tower jettisoned during the boost
phase of flight. See figure 9-6 for vehicle configuration and missiontime lines.
SPACECRAFT SEPARATION.
The CSM separates from the SIVB/IU/SLA at vehicle station
2033.799 (reference Section I, Vehicle Profile) approximately 6 min
after orbit insertion. Explosive fuse assemblies sever the CSM from
the SLA, and sever the adapter longitudinally into four panelsections.
The panels then rotate about hinges attached to the payload adapter
with each panel deploying to a 45-deg open position, thus comple-
ting the spacecraft separation. (This is true of one Skylab Saturn
IB launch yehicle; the others have jettisonable SLA panels same
as the Apollo Saturn V launch vehicles.) After separation, the
S-IVB/IU/SLA executes maneuvers required to maintain the vehi-
cle in the proper attitude for performing the Thermal Control
Coating Experiment (M-415).
S-IVB ACTIVE LIFETIME. S-IVB ORBITAL DECAY HISTORY "
The S-1VB stage as initially designed has an orbital coast capability
of 4.5 hr. This capability has been extended for the SL-2 and SL-4
missions through a number of changes to improve the temperature
control of certain components and assemblies. To accommodate
the Skylab experiments, these two stages now have a lifetime of
approximately 7.5 hr.
APS propellant depletion occurred on SA-205 between stationcontacts at 15 hr 30 min and 16 hr 20 min. Data indicated that
attitude control was normal prior to APS propellant depletion.
ORBITAL CONFIGURATION.
ORBITAL INSERTION.
At orbftal insertion 9 min 51.9 sec into flight (predicted to be 10
(NM)140r-
120F-
I001-
801-
601-
401-
201-
0 =-
(KM)
50
APOGEE
xJ
PERIGEE_0
)0 ---
0
LAUNCH DATE: MAy 1, 1973
DENSITY: +2=' NOMI
I16 24
TIME (HR)
-2a
32
Figure 9-5
C-H 20148
9-5
Section IX Mission Variables and Constraints
LAUNCH VEHICLE CONFIGURATION AND MISSION TIMELINES
CSM/S-IVB
SEPARATION -_ ..
SEE TABLE Fllllllllllllln _11111_ IIII
APOLLO /"CONFIGURATION
AFTER
SEPARATIONFROM S-IVB --.,/
SATURN IB
LAUNCH VEHICLE
CONFIGURATION
AFTER S-IB/S-IVBSEPARATION
i=_..,_.,_SEPARATION
RN IB
LAUNCH VEHICLE
CONFIGURATION
AT LIFTOFF
TYPICAL MISSION TIMELINES
INITIATE
00:00:00
00: 02:22
00:02:45.6
00:09:41.9
00: 09:51.9
00: 09: 42. I
00:10:02. t
00:14:. 42. I
00:55:34.1
FLIGHT TIME
(HR: MI N: SEC)
END
00:10:02:1
00:14: 42. |
DESCRIPTION
LIFTOFF (FIRST MOTION)
S-I B/S-IVB SEPARATION
LAUNCH ESCAPE TOWER JETTISON
S-IVB ENGINE CUTOFF
ORBITAL INSERTION
MAI NTAI N CUTOFF INERTIAL ATTITUDE FOR 20
SECONDS AFTER INITIATION OF TIME BASE
FOUR CTB4).
I NITIATE A PITCH MANEUVER TO ALIGN THE
S-IVB/CSM ALONG THE LOCAL HORIZONTAL,
NOSE LEADING, PISITION I DOWN. MAINTAIN
ORBITAL RATE.
NOMI NAL CSM SEPARATION.
INITIATE THERMAL COATING EXPERIMENT M415.
LIFTOFF
Figure 9-6
C-H 14552-2
9-6
APPENDIX A
ABBREVIATIONS, SIGNS, AND SYMBOLS
A
ABMA
ac
accel
ACE
ACM
AFETR
alt
AM
amb
AN
AOA
approxAPS
ARPA
ASAP
ASC
ASD
ASI
assyASTM
ATM
atm
att
auto
aux
avgAVP
BEF
bhpBMAG
bpsBSE
BTU
calib
calipsCAPCOM
CASTS
CAT
C/B
CCW
C/D
CDC
CDDT
CDFCDSC
CDU
CEI
ABBREVIATIONS.
A
ampere
Army Ballistic Missile Agency
alternating currentaccelerometer
automatic checkout equipmentactuation control module
Air Force Eastern Test Rangealtitude
Airlock Module
ambient
ascending node
angle of attack
approximately
auxiliary propulsion systemAdvance Research Projects Agency
auxiliary storage and playback
accelerometer signal conditioner
abort summary document
augmented spark igniter
assemblyAmerican Society for Testing
Materials
Apollo Telescope Mount
atmosphereattenuator or attitude
automatic
auxiliary
averageaddress verification pulse
B
blunt end forward
brake horsepower
body-mounted attitude gyro
bits per second
booster systems engineerBritish thermal unit
C
calibration
calibrational pressure switch
spacecraft communicatorCountdown and Status Transmission
Systems
control attenuating timercircuit breaker
counterclockwise
collect/disperse (computer)countdown clock
countdown demonstration test
confined detonating fuseCommunication Distribution and
Switching Center
coupling data unitcontract end item
CG center of gravitychan channel
char characteristics
CIF Central Instrumentation Facility
CIU computer interface unit
CKAFS Cape Kennedy Air Force Stationcm centimeter
CM command module
CMC command module computer
cmd commandCMR command module receiver
COD cross-over detectors
cont control
convtr converter
CRES corrosion resistant (steel)
CRP computer reset pulseCSM command service module
CSP control signal processorCST coast
CT components test
C-T crawler-transporterCW clockwise
cx/ct control transmitter/control trans-
former
D
db decibel
DB disagreement bitdc direct current
DCS digital command systemdev deviation
DDAS digital data acquisition system
decr decreasing
DEE digital events evaluator
deg degree, angulardest destruct
DI discrete inputdia diameter
dir directional
distr distributor
DN descending node
DO discrete outputDOD Department of Defense
DPF dynamic pressure feedback
DRS data receiving station
dyn dynamic
E
EBW exploding bridgewireECS environmental control system
EDS emergency detection systemelec electrical
ELS earth landing system
EMR engine mixture ratio
EMS entry monitor system
eng engine
equip equipment
A-1
!
Appendix A
err
ESE
ETD
ETR
FABU
FCC
FCSMFCVB
FDAI
FIDO
FLSC
fit
FM
F/M
FOD
FOMR
FPR
FRR
FRT
ft
fwd
gG
galGCS
GDC
genGG
G MTC
G&N
gndGOX
gPf
gpm
grGRRGSCU
GSE
GSFC
guid
HEP
HGA
HOSC
H.P.
hphr
H/W
hydHz
H-I
H-lC
H-1D
IA
ICC
ID
IECO
IF
A-2
error
electrical support equipment
end thrust decay
Eastern Test Range
F
fuel additive blender unit
flight control computerflight combustion stability monitorflow control valve box
flight director attitude indicator
flight dynamics officer
flexible linear-shaped charge
flight
frequency modulation
thrust acceleration (force/mass)
Flight Operations Director
Flight Operations Management Room
flight performance reserve
flight readiness reportflight readiness testfeet or foot
forward
gramgravitational constant
gallon
guidance cutoff signal
gyro display coupler
generator
gas generatorGreenwich mean time clock
guidance and navigation
ground
gaseous oxygengrains per foot
gallons per minute
grain
guidance reference release
ground support cooling unit
ground support equipment
Goddard Space Flight Center
guidance
H
hardware evaluation program
hazardous gas analyzer
Huntsville Operations SupportCenter
high pressure
horsepowerhour
hardwire
hydraulichertz
Saturn IB first stage (S-IB) engine
S-IB stage inboard engine
S-IB stage outboard engine
I
input axisInterstate Commerce Commission
identification or inside diameter
inboard engine cutoff
intermediate frequency
IFV
IGM
IMCC
impIMU
IMV
in.
inbd
incr
ind
instl
instr
int
invtr
IOA
IODC
IOR
IOS
IP
IU
J-2
jett
k
KSC
lb
lbf
Ibm
LCC
LE
LES
LET
LIEF
LLS
LM
lox
L.P.
LPGG
LSC
LUT
LV, L/V
LVDA
LVDC
LVO
m
M
man
max
MCC
MD
MDA
MDC
MDF
meas
med
MESC
MFCO
igniter fuel valve
iterative guidance mode
Integrated Mission Control Center
impulseinertial measurement unit
Ignition monitor valveinch
inboard
increasingindication
installation
instrumentation
mternal
inverter
input/output address
input/output data channel
input/output register (buffer)
input/output sense
impact predictionInstrument Unit
Saturn IB second stage (S-IVB)
engine
jettison
K
kilo (prefix)
Kennedy Space Center
L
pound
pound (force)
pound (mass)launch control center
launch escape
launch escape system
launch escape tower
Launch Information Exchange
Facility
liquid level sensorlunar module
liquid oxygen
low pressure
liquid propellant gas generator
linear shaped chargelauncher umbilical tower
launch vehicle
launch vehicle data adapter
launch vehicle digital computer
launch vehicle operations
M
milli (prefix) or meter
mega (prefix)manual
maximum
Mission Control Center
mission director
Multiple Docking Adapter
main display console
mild detonating fusemeasurement
medium
master events sequence controllermanual fuel cutoff
MFCVMFVMGSEmiMILAminmlMLMLVMMHmo
MOCR
mod
MOV
MRCV
ms
MSC
MSFC
MSO
MSS
MTVC
MUX
N
N/A
NASA
NC
negN.G.& C
NM
No.
NO
NP
NPSH
NPV
OA
OAT
OECO
OETD
OIS
ord
OSC
OSR
OWS
oxid
OZ
PAM
PC
PCD
PCM
PDS
PEA
PETN
PIF
pkg
pneupos
modulating flow control valvemain fuel valve
mechanical ground support equipmile
Merritt Island Launch Activityminute or minimum
milliliter
mobile launcher
main lox valve
monomethylhydrazinemonth
Mission Operations Control Roommodule or model
main oxidizer value
mixture ratio control valvemillisecond
Manned Spacecraft Center
Marshall Space Flight Center
Mission Support Operationsmobile service structure
manual thrust vector control
multiplexer
N
Newton
not applicable or not available
National Aeronautics and SpaceAdministration
normally closed
negative
navigation, guidance, & controlnautical mile
number
normally open
north pole
net pressure suction head
non-propulsive vent
O
output axisoverall test
outboard engine cutoff
outboard engine thrust decay
Operations Intercommunications
Systemordnanceoscillator
Operations Support Room
Orbital Workshopoxidizer
ounce (torque)
P
pulse amplitude modulation
pitch control
pneumatic control distributor
pulse code modulation
propellant dispersion
system
platform electronics
assembly
pentaerythritol tetranitratePublic Information Facility
package
pneumatics
positive
POS
POST
ppm
prep
press
propprplnt
psi
psia
psid
psig
PTCS
PU
pwr
O
Oa
Q-D
RACS
RAD (rad)
RASM
RC
R. CAL
RCC
RCS
RCVR
R&D
RDM
RDSM
RDX
ref
regrevRF
RFI
RNG
rpmRP-I
RS
RSCR
RSO
rss
RTCC
SA
S&A
SIC
scfm
Appendix A
position
priority of systems tests
parts per million
preparation
pressure
propellant
propellant
pounds per square inch
pounds per square inch
(atmospheric)
pounds per square inch
(differential)
pounds per square inch
(gage)
propellant tanking computer
system
propellant utilization
power
Q
dynamic pressure
angle-of-attack/dynamic pressure product
quick disconnect
R
remote automatic cali-
bration system
a circle segment equal to
180 deg/rrremote analog
submultiplexerresistor capacitorradiation calorimeter
Recovery Command andControl Center
reaction control systemreceiver
research and development
remote digital multiplexer
remote digital
submultiplexer
cyclotrimethylenetrinitramine
reference
regulatorrevolution
radio frequency
radio frequencyinterference
rangingrevolutions per minute
rocket propellant grade 1(fuel)
range safety
range safety commandreceiver
range safety officer
root-sum-square
Real-Time Computer
Complex
S
Saturn or service arm
sating and arming
spacecraft
standard cubic feet perminute
A-3
Appendix A
scim
SCT
sec
SECO
sel
sep
seqsf
S-IB
sigSIR
S-IVB
SL
SLA
SLDS
SM
SMC
SOCR
sol
spSPAN
SPGG
SPS
SRA
SSR
sta
stbystd
STDN
STDV
STP
sw
SWS
sys
TB
TC
standard cubic inches perminute
scanning telescopesecond
S-IVB engine cutoffselector
separation
sequencescaling factor
first stage of Saturn IBlaunch vehicle
signal
systems integration rack
second stage of Saturn IBlaunch vehicle
Skylab
spacecraft lunar (module)
adapter
Skylab Launch Data Systemservice, module
steering misalignmentcorrections
Sustained OperationsControl Room
solenoid
specific
spacecraft analysis
solid propellant gas
generator
service propulsion system
spin reference axis
Staff Support Roomstation
standbystandard
Spacecraft Tracking and Data Network
start tank discharge valvestandard pressureswitch
Saturn Workshop
system
T
time base
thrust chamber
TCS
tempTER
TLC
TLI
TM
TMR
TOPS
turb
TV
TVC
twr
typ
UDMH
UHF
umb
U/R
USB
V
VAB
Vac
VCS
Vdc
VHF
VppVSWR
W
wt
XFER
XMTR
thermal conditioning
system or terminal
countdown sequencer
temperature
telemetry executiveroutine
simultaneous memoryerror
translunar injection
telemetry
triple modular redundancythrust ok pressure
switch
turbine
television
thrust vector control
tower
typical
U
unsymmetrical
dimethylhydrazine
ultra-high frequencyumbilical
ullage rocketunified S-band
V
volt or velocity
vehicle assembly building
volts alternating currentVoice Communications
Systemvolts direct current
very-high frequency
volts peak-to-peak
voltage standing wave ratio
W
watt
weight
X
transfer
transmitter
A-4
@°APi
°C
o F
o R
cLGHe
GH 2
GN2He
H_OKOH
LH 2
N204
qZA
>___<>
<
f_
B,
0
O"
¥
at
degrees (scale) AmericanPetroleum Institute
degrees Celcius
degrees Fahrenheit
degrees Rankincenterline
gaseous helium
gaseous hydrogen
gaseous nitrogenhelium
water
potassium hydroxide
liquid hydrogen
nitrogen tetroxide
aerodynamic pressureangle
change
greater than or equal toless than or equal to
greater thanless than
difference
nearly equal todefined as
ohm
alpha, angle of attack
beta, feedback signals from
H-I engine actuators
steering command output
gamma, lateralacceleration
theta, platform gimbal
angle
platform gimbal angle(vehicle attitude)
lambda, longitude
mu, micro (prefix)
sigma, an occurrence pro-
bability symbol express-
ing a percentage of all
possible values in a
given parameter.
tau, control accelerometer
signals
phi, attitude rate
SIGNS AND SYMBOLS.
LX
X,X_Xy
×,Xy×,
DX
¢¢,¢_¢,
eTAi
PRY
UVW
x
XYZ
XIYIZI
x Yz
X YZ
X4Y4Z4
system
DD@
---0
launch site latitude
chi, desired attitude change
guidance command angles(desired vehicle attitude)
guidance command angles
(Euler angles)thrust direction definition
average change for each
minor loop guidance
computation
psi, attitude errorattitude error signals
(steering commands)
total range angleazimuth
inclination of orbit
pitch, roll, and yaw axes
gravitational coordinate
system
inertial velocity
incremental velocity
platform gimbal pivot axesmeasurement coordinate
systeminertial coordinate system
injection coordinate
system
Logic Symbols.
AND gate
OR gate
time delay in sec
signal, presence of
signal, absence of
Appendix A
A-5